低轨道航空器辐射环境和表面充电效应研究
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摘要
低地球轨道是目前航天器的主要应用轨道之一。该轨道区域中高能粒子辐射环境和等离子体环境是诱发低轨道航天器异常故障的主要空间环境因素,研究其变化规律和影响航天器的效应,可以为今后航天器的空间环境防护设计提供技术基础,在空间环境研究和航天工程应用两方面都有重要意义。本文针对低轨道辐射环境,开展了低高度内辐射带高能质子环境研究,分析了太阳周期活动对低高度内辐射带高能质子环境的影响;另外针对低轨道等离子体环境的效应,开展了航天器表面充电和高电压太阳电池阵电流泄漏效应研究。
在低高度内辐射带高能质子环境研究中,本文利用NOAA-15卫星1998年到2011年近13年的高能质子全向通量的观测资料,分析了一个太阳活动周内低高度内辐射带高能质子通量的分布变化特性及其物理原因,比较了观测结果与AP8模型的不同。研究表明,低高度内辐射带高能质子通量与太阳活动水平的反相关关系与磁壳参数??值、磁场??值有关;??值越低、??值越大的空间点,其高能质子通量与太阳活动水平的反向相关性越明显。高能质子通量随太阳活动水平的变化存在明显的滞后现象,??值越高、??值越小的空间点,滞后现象就越明显,滞后严重时可以达到一年左右的时间。高能质子通量随太阳活动水平变化的滞后现象反映低高度内辐射带高能质子的源与损失达到平衡是一个中长期过程,这种不平衡性导致相同太阳活动水平下,同一点的高能质子通量也可能存在一定差异。通过与AP8模型计算结果的比较,发现在考虑磁场长期变化的情况下,对于高磁场值处的高能质子通量,模型计算结果明显高于卫星观测资料。我们分析认为,利用AP8模型时,仅考虑磁场长期变化对质子通量的影响可能会夸大低高度内辐射带局部高能质子通量的增强。
在低轨道航天器表面充电分析研究中,本文在简化的球形几何模型下,基于表面电流平衡方程,建立了一个一维条件下的航天器表面充电评估计算方法。利用该方法编制的计算程序,分析研究了低轨道航天器表面充电与等离子体环境、表面材料特性等的关系。分析研究表明,在低轨道等离子体环境中,高表面充电与高温沉降电子的总通量无关,而与沉降粒子中能量大于十几keV以上的电子通量有关,能量大于10keV以上的电子通量小于108 cm?2s?1sr?1时,很难发生表面高充电;背景电离层等离子体会制约表面充电,其密度越高,航天器表面充电电位越低,密度高于一定限度时,不会发生表面高充电。另外,光照条件会极大地抑制表面高充电。表面入射电子的最大二次电子发射率是影响表面充电的最重要的材料特性参数,电子的最大二次电子发射率大于3时,表面很难充高电位。当航天器表面存在多种材料时,各种材料的充电水平会互相影响,表面材料间的导电特性可以决定这种影响的程度;若航天器表面材料之间绝缘或导通性不好,尽管航天器整体采用了抗表面充电的材料,局部其它材料表面仍可能独立出现高充电;若航天器表面材料均为导体,局部表面材料使用或处理不当,会影响整体表面的充电水平。
在低轨道航天器高电压太阳电池阵电流泄漏效应研究中,本文在一种简化的太阳电池阵结构模型的基础上,利用地面实验数据得到的太阳电池阵电流收集经验模式,基于电流平衡,建立了一种计算航天器高电压太阳电池阵电流泄漏的方法。利用该方法我们分析计算了高电压太阳电池阵泄漏电流损耗与低轨道等离子体环境、电池阵电压和电池阵裸露金属面积之间的关系。分析研究表明,影响电流泄漏的主要环境因素是等离子体密度。等离子体密度越高,影响越大。电流泄漏影响最严重的区域是电离层等离子体密度最高的300?400km高度的轨道区域。电流泄漏引起的功率损失与太阳电池阵电压呈指数关系,电压越高,功率损失越大。对于200V以下的电池阵,功率损耗较小,远低于电源系统总功率的1%;当电压达到1000V,功率损耗可达到总功率的2%左右。电流泄漏与太阳电池阵的裸露金属导体表面的面积呈正比关系,因此通过减少太阳电池阵裸露面积,可以降低电流泄漏的影响。
Low Earth Orbit(LEO) is one of the most important orbits for spacecraft atpresent. In LEO region, the high energetic particle and plasma are the major spaceenvironment factors which could cause spacecraft system failures and anomalies.Investigating the characteristics of space environment and effects to spacecraft canprovide the technical basis for spacecraft design to defend the space environmentthreat, and is also benefit to space environment research. In this dissertation, thefeature about high energy proton environment of inner radiation belt in the lowaltitude region is investigated at first. Then the research about plasma effects onLEO spacecraft are carried out, which include the analysis of spacecraft surfacecharging and the simulation of the current leakage effect.
The NOAA-15 high energy proton observation from 1998 to 2011 is used toanalyze the effect of solar cycle activity on high energy proton flux. The statisticresearch indicates that there is an inverse correlative relationship between theproton flux in inner radiation belt and solar activity. This anti-correlation is relatedto geomagnetic coordinates ?? and ??, and more significant with the increasing of?? and decreasing of ??. There is also a phase lag between the solar activity andthe proton flux. This hysteresis effect is more obvious in the region with smaller?? or larger ??. The lag can reach one year in some region. This hysteresis effectmeans it takes a long time to reach the dynamic balance between the source andthe loss for the proton of inner radiation belt in the low altitude region. Theunbalance between the source and loss is the reason why the intensity of protonflux at the same solar activity is different. The comparison with the result of AP8model indicates the energetic proton flux from AP8 is higher than the satellite’sobservation in the region with large B, which suggests the Ap8 model may overstatethe proton flux enhancement at inner radiation belt in the low altitude region ifonly the long-term variation of magnetic field considered .
The technique for evaluating spacecraft surface charging is developed in onedimension condition with the geometric model of spacecraft simplified to a sphere.This method is based on the application of the current balance equation and empirical formula of secondary current. The variation of surface charging lev-els with plasma conditions in LEO and material properties was studied by usingthis method. The result shows that the spacecraft charging level has little cor-relation with precipitation electron total flux while it obviously increases withelectron fluxes above 10KeV. When the electron fluxes above 10KeV below 108cm?2s?1sr?1, the spacecraft can be hardly charged to high level (several tens ofVolts to thousands of Volts). The background thermal plasma restrains the surfacecharging level. The surface charging level decreases with thermal plasma density,and there is a threshold value of plasma density beyond which high level surfacecharging can not occur. The sunlight can hinder surface charging. The mostimportant material property affecting the surface charging is the maximum sec-ondary electron emission yield for electrons. When the emission yield is beyond 3,the spacecraft can not be charged high level. If there are more than two materialsover the surface of spacecraft, the charging status will be influenced by each otherand the charging result is determined by the conductivity among surface materials.When it is insulated among surface materials or the conductivity are not good, itis still possible for the high level charging in some small local area although themost surface of spacecraft adopted the resisting charging materials. When all thematerials over the surface are conductive, the whole spacecraft charging level canstill be affected if the regional surface materials processing is inappropriate.
The last part of this dissertation is the development for calculating currentleakage effect of high voltage solar array in LEO. The method is based on currentbalance, a simplified solar array structure model and an empirical current collectionmodel of solar array. Using this method, the relationships between power losscaused by current leakage and plasma environment, solar array voltage, the areaof exposed conductor are analyzed. The research shows the power loss of solararray declines quickly with the increasing of orbital altitude. The current leakagein the region among 300-400 kilometers is most serious because of the high plasmadensity. The power loss increases exponentially with the voltage of solar array.Meanwhile, we also noticed that the power loss is far less than 1% of total poweras the voltage of solar array is below 200V. The power loss of the solar arrayincreases linearly with the area of exposed conductor. Based on this result, the current leakage effect can be reduced by decreasing the conductor’s exposure areaof solar array.
在低高度内辐射带高能质子环境研究中,本文利用NOAA-15卫星1998年到2011年近13年的高能质子全向通量的观测资料,分析了一个太阳活动周内低高度内辐射带高能质子通量的分布变化特性及其物理原因,比较了观测结果与AP8模型的不同。研究表明,低高度内辐射带高能质子通量与太阳活动水平的反相关关系与磁壳参数??值、磁场??值有关;??值越低、??值越大的空间点,其高能质子通量与太阳活动水平的反向相关性越明显。高能质子通量随太阳活动水平的变化存在明显的滞后现象,??值越高、??值越小的空间点,滞后现象就越明显,滞后严重时可以达到一年左右的时间。高能质子通量随太阳活动水平变化的滞后现象反映低高度内辐射带高能质子的源与损失达到平衡是一个中长期过程,这种不平衡性导致相同太阳活动水平下,同一点的高能质子通量也可能存在一定差异。通过与AP8模型计算结果的比较,发现在考虑磁场长期变化的情况下,对于高磁场值处的高能质子通量,模型计算结果明显高于卫星观测资料。我们分析认为,利用AP8模型时,仅考虑磁场长期变化对质子通量的影响可能会夸大低高度内辐射带局部高能质子通量的增强。
在低轨道航天器表面充电分析研究中,本文在简化的球形几何模型下,基于表面电流平衡方程,建立了一个一维条件下的航天器表面充电评估计算方法。利用该方法编制的计算程序,分析研究了低轨道航天器表面充电与等离子体环境、表面材料特性等的关系。分析研究表明,在低轨道等离子体环境中,高表面充电与高温沉降电子的总通量无关,而与沉降粒子中能量大于十几keV以上的电子通量有关,能量大于10keV以上的电子通量小于108 cm?2s?1sr?1时,很难发生表面高充电;背景电离层等离子体会制约表面充电,其密度越高,航天器表面充电电位越低,密度高于一定限度时,不会发生表面高充电。另外,光照条件会极大地抑制表面高充电。表面入射电子的最大二次电子发射率是影响表面充电的最重要的材料特性参数,电子的最大二次电子发射率大于3时,表面很难充高电位。当航天器表面存在多种材料时,各种材料的充电水平会互相影响,表面材料间的导电特性可以决定这种影响的程度;若航天器表面材料之间绝缘或导通性不好,尽管航天器整体采用了抗表面充电的材料,局部其它材料表面仍可能独立出现高充电;若航天器表面材料均为导体,局部表面材料使用或处理不当,会影响整体表面的充电水平。
在低轨道航天器高电压太阳电池阵电流泄漏效应研究中,本文在一种简化的太阳电池阵结构模型的基础上,利用地面实验数据得到的太阳电池阵电流收集经验模式,基于电流平衡,建立了一种计算航天器高电压太阳电池阵电流泄漏的方法。利用该方法我们分析计算了高电压太阳电池阵泄漏电流损耗与低轨道等离子体环境、电池阵电压和电池阵裸露金属面积之间的关系。分析研究表明,影响电流泄漏的主要环境因素是等离子体密度。等离子体密度越高,影响越大。电流泄漏影响最严重的区域是电离层等离子体密度最高的300?400km高度的轨道区域。电流泄漏引起的功率损失与太阳电池阵电压呈指数关系,电压越高,功率损失越大。对于200V以下的电池阵,功率损耗较小,远低于电源系统总功率的1%;当电压达到1000V,功率损耗可达到总功率的2%左右。电流泄漏与太阳电池阵的裸露金属导体表面的面积呈正比关系,因此通过减少太阳电池阵裸露面积,可以降低电流泄漏的影响。
Low Earth Orbit(LEO) is one of the most important orbits for spacecraft atpresent. In LEO region, the high energetic particle and plasma are the major spaceenvironment factors which could cause spacecraft system failures and anomalies.Investigating the characteristics of space environment and effects to spacecraft canprovide the technical basis for spacecraft design to defend the space environmentthreat, and is also benefit to space environment research. In this dissertation, thefeature about high energy proton environment of inner radiation belt in the lowaltitude region is investigated at first. Then the research about plasma effects onLEO spacecraft are carried out, which include the analysis of spacecraft surfacecharging and the simulation of the current leakage effect.
The NOAA-15 high energy proton observation from 1998 to 2011 is used toanalyze the effect of solar cycle activity on high energy proton flux. The statisticresearch indicates that there is an inverse correlative relationship between theproton flux in inner radiation belt and solar activity. This anti-correlation is relatedto geomagnetic coordinates ?? and ??, and more significant with the increasing of?? and decreasing of ??. There is also a phase lag between the solar activity andthe proton flux. This hysteresis effect is more obvious in the region with smaller?? or larger ??. The lag can reach one year in some region. This hysteresis effectmeans it takes a long time to reach the dynamic balance between the source andthe loss for the proton of inner radiation belt in the low altitude region. Theunbalance between the source and loss is the reason why the intensity of protonflux at the same solar activity is different. The comparison with the result of AP8model indicates the energetic proton flux from AP8 is higher than the satellite’sobservation in the region with large B, which suggests the Ap8 model may overstatethe proton flux enhancement at inner radiation belt in the low altitude region ifonly the long-term variation of magnetic field considered .
The technique for evaluating spacecraft surface charging is developed in onedimension condition with the geometric model of spacecraft simplified to a sphere.This method is based on the application of the current balance equation and empirical formula of secondary current. The variation of surface charging lev-els with plasma conditions in LEO and material properties was studied by usingthis method. The result shows that the spacecraft charging level has little cor-relation with precipitation electron total flux while it obviously increases withelectron fluxes above 10KeV. When the electron fluxes above 10KeV below 108cm?2s?1sr?1, the spacecraft can be hardly charged to high level (several tens ofVolts to thousands of Volts). The background thermal plasma restrains the surfacecharging level. The surface charging level decreases with thermal plasma density,and there is a threshold value of plasma density beyond which high level surfacecharging can not occur. The sunlight can hinder surface charging. The mostimportant material property affecting the surface charging is the maximum sec-ondary electron emission yield for electrons. When the emission yield is beyond 3,the spacecraft can not be charged high level. If there are more than two materialsover the surface of spacecraft, the charging status will be influenced by each otherand the charging result is determined by the conductivity among surface materials.When it is insulated among surface materials or the conductivity are not good, itis still possible for the high level charging in some small local area although themost surface of spacecraft adopted the resisting charging materials. When all thematerials over the surface are conductive, the whole spacecraft charging level canstill be affected if the regional surface materials processing is inappropriate.
The last part of this dissertation is the development for calculating currentleakage effect of high voltage solar array in LEO. The method is based on currentbalance, a simplified solar array structure model and an empirical current collectionmodel of solar array. Using this method, the relationships between power losscaused by current leakage and plasma environment, solar array voltage, the areaof exposed conductor are analyzed. The research shows the power loss of solararray declines quickly with the increasing of orbital altitude. The current leakagein the region among 300-400 kilometers is most serious because of the high plasmadensity. The power loss increases exponentially with the voltage of solar array.Meanwhile, we also noticed that the power loss is far less than 1% of total poweras the voltage of solar array is below 200V. The power loss of the solar arrayincreases linearly with the area of exposed conductor. Based on this result, the current leakage effect can be reduced by decreasing the conductor’s exposure areaof solar array.
引文
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叶宗海,陈贵福. 1993.在“风云一号(B)”卫星上空间粒子环境的探测结果[J].中国科学A辑, 23(005):538–544.
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