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为验证热设计的正确性,对微型星敏的热控方案进行了热分析。在I-DEAS中建立其热分析模型如图3所示。其模型选取壳单元类型,划分了2 245个壳单元,采用设置热耦合的方式简化结构。
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外热流分析是热设计的基础。对外热流分析而言,影响最大的热源来自太阳的直接照射热流。外热流的变化是导致星敏感器组件温度波动的主要原因,微型星敏外热流随着卫星进出阴影区发生巨大的变化,使得星敏感器本体乃至安装面产生温度波动。其中,高低温工况外热流如图4和图5所示。
地球轨道高度600 km,低温工况与高温工况参数选取定义如表1所示,表中β为太阳光与轨道面的夹角,假设地球漫反射体,反照率取平均反照率,取常数。
Case Solar constant/
W·m−2β/(°) Internal heat source
power/WHigh temperature case 1414 73 5 Low temperature case 1309 0 5 Table 1. Working condition parameters
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微型星敏在未采取热控措施的情况下,微型星敏本体的温度随时间变化分布曲线如图6所示。由分析结果可以看出,未采用合理的热控措施微型星敏本体高温工况的温度为61.1 ℃,低温工况下的最低温度−26.3 ℃,超出温度指标−20~+40 ℃,既有在高温工况下的散热需求,也有低温工况下的加热需求。
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微型星敏在实施热控措施后,微型星敏本体的温度随时间变化分布曲线如图7所示。仿真结果表明,采用合理的热控措施后,微型星敏本体高温工况的温度范围为15.0~19.0 ℃,在轨道周期瞬态情况下微型星敏本体的温度随时间的波动范围是4 ℃;低温工况下的温度范围−4.9~5.1 ℃,在轨道周期瞬态情况下微型星敏本体的温度随时间的波动范围是10 ℃以内,满足热控指标要求。
图中,高温工况热控温度范围明显优于低温工况的热控温度范围,主要由于在低温工况下,星敏温度一旦低于控温阈值下限,根据控温策略加热回路迅速启动补偿星敏本体温度,当星敏本体温度到达控温阈值上限后,加热回路关闭;而在高温工况加热回路没有启动,星敏本体温度变化受外热流变化等因素影响,因此,在低温工况、高温工况,导致了星敏本体温度曲线的不同表现。
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微型星敏本体与遮光罩间在采取粘贴石墨导热带措施的情况后,开展星敏本体与遮光罩温度均匀性仿真,结果如图10所示。经结果分析表明,微型星敏本体与遮光罩之间温度波动范围小,说明石墨导热带措施确保了微型星敏本体与遮光罩之间星敏整体良好的均温性。
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为考核星敏感器组件热设计的正确性,微型星敏感器及其支架参与完成了整器平台的热平衡试验。星敏感器组件与整器平台一起进行热平衡试验,保证了边界的正确性。热试验条件如下:真空室内真空度优于1.3×10−3 Pa;热沉温度≤100 K。
热试验中,微型星敏的热控设计状态与前面相同。采用红外加热笼模拟外热流。但由于星敏支架形状复杂,各部位的外热流密度存在较大差异。
试验结果见表2,温度数据与仿真分析中结果一致性良好,星敏感器组件的热设计能够满足热控指标。
High temperature case Low temperature case Maximum temperature Minimum temperature Maximum temperature Minimum temperature 20.61 14.47 6.49 −4.98 Table 2. Thermal balance test data results (Unit: ℃)
Thermal control design and simulation of micro star sensor
doi: 10.3788/IRLA20220116
- Received Date: 2022-04-10
- Rev Recd Date: 2022-06-10
- Publish Date: 2022-11-30
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Key words:
- thermal control design /
- thermal analysis /
- micro star sensor
Abstract: In order to ensure the normal operation of the star sensor during the on-orbit mission of the application platform, it needs to be thermally designed. First, a process of thermal analysis and optimization was proposed, which combined with the conditions of the external heat flow, installation layout and working mode of the micro star sensor assembly. In the process of thermal analysis optimization, the influence of various factors, such as optical, mechanical and thermal, was considered. Second, a thermal control scheme of the micro star sensor assembly was designed. The thermal control scheme proposed a design idea of using active electric heating and homogenizing the temperature between the light-shield and the star sensor body, which solved the problems of complex space thermal environment, higher temperature control requirements, and heat dissipation path limited by a installation structure during the on-orbit period of the micro star sensor module. This scheme ensured the effective and reliable work of the micro star sensor assembly. Third, an I-DEAS/TMG finite element mode was established. The thermal control simulation of the micro star sensor assembly under high and low temperature conditions was carried out, and the simulation results of the temperature distribution and uniformity of the star sensor assembly were analyzed. Finally, a ground test was carried out to ensure the correctness of the thermal control scheme, and the test results met the thermal requirement of the star sensor assembly. This paper can provide a reference for the following thermal design of micro star sensor assembly of on-orbit platforms.