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监测仪包络尺寸(X×Y×Z):1 070 mm×930 mm×650.8 mm,由底板、外罩、遮光罩、光学镜头组件及各种信息处理器和控制器等11种电子设备组成,布局图(图1),电子设备E,光学镜头O,电机M。电子设备3~6分别安装在镜头2~5镜筒上,其余设备和镜头都安装在底板上。监测仪在卫星上的位置(图2)。表1为内热源热耗和工作时间,卫星轨道周期为102 min。
镜7为二维光学转动机构,镜7 CT轴布置1台步进电机,AT轴布置1台音圈电机。镜7工作时,AT轴音圈电机跟随镜框在CT轴步进电机带动下绕CT轴转动,同时,音圈电机转子转动带动镜体绕AT轴俯仰,音圈电机热耗集中在其转子线圈上,是二维转动热源。镜7组件及音圈电机结构示意图见图3。
表 1 内热源热耗及工作时间
Table 1. Heat consumption and operating time of calorigenic equipments
Calorigenic
equipmentsStand by
mode/WObservation
mode/WCalibration
mode/WOperating
time on
orbit/minElectric device 1 11 11 11 102 Electric device 2 0 3 3 68 Electric device 3 0 0.3 0.3 63 Electric device 4 0.9 1.2 1.2 102 Electric device 5 0.9 1.2 1.2 102 Electric device 6 0.9 1.2 1.2 102 Electric device 7 0 0 4 3 Electric device 8 0 20 20 63 Electric device 9 0 0 21.1 3 Electric device 10 8 8 8 102 Electric device 11 0 4 4 52 Voice coil motor 0 2 2 52 Stepper motor 0 1.7 1.7 52 监测仪各部组件控温要求如表2所示。镜7组件位于入光口处,其镜体控温指标分为工作时和非工作时两种,非工作时指标要求相对宽松。电子设备3~6热耗集中在壳体内部的发热部件热端,发热部件热端靠耳片固定在壳体上,热端温度要求控制在0~20 ℃。
表 2 气体监测仪各部组件控温要求
Table 2. Temperature control requirements of gas monitoring sensor components
Components Temperature requirement/℃ Operating time Non-operating time Optical lens 7 19±3 19±7 Optical lens 1–6,8 20±2 Voice coil motor of optical lens 7 0–85 Stepper motor of optical lens 7 –10–80 Electric device 1,2,7–11 –10–45 Electric device 3–6 0–20 -
通过热平衡试验,获取在轨热边界条件下监测仪各部组件温度数据,对监测仪热设计进行验证。监测仪在轨热边界条件及模拟方法如表3所示。监测仪各部组件按照设计状态进行热控实施和总装。试验工况设置了低温初期瞬态工况和高温末期瞬态工况两种,如表4所示。
表 3 气体监测仪在轨热边界条件及试验模拟方法
Table 3. Thermal boundary conditions of gas monitoring sensor on orbit and simulation methods of these conditions during ground test
Thermal boundary conditions Simulation methods during ground test Vacuum and space cold black background Use space environment simulator Thermal environment on satellite platform Design simulation tool of satellite platform,which is shown in Fig.9, Fig.5
Control the temperature of simulation tool according to the given temperature boundary of satellite;
Cover simulation tool surfaces facing gas monitoring sensor with MLI and put heating circuit on surface of the MLI to obtain orbit heat flux absorbed by MLIOrbit heat flux Use infrared heating cage and flux sensor to abtain heat flux incidenting to the entrance of earth baffle;
Put heating circuit on MLI surfaces and radiator back surface to obtain orbit heat flux absorbed by themHeat consumption of calorigenic equipments Calorigenic equipments of gas monitoring sensor operate during test according to the normal working mode on orbit 表 4 试验工况
Table 4. Operating conditions of test
Case Orbit heat flux One orbit working mode Temperature boundary
of satellite/℃Cold case Minimum heat flux throughout the life cycle Standby mode −5 Hot case Maximum heat flux throughout the life cycle Standby mode→calibration mode→
observation mode→standby mode45 -
图10给出了低温初期瞬态工况和高温末期瞬态工况监测仪镜头组件连续四轨的温度曲线,表5对试验结果进行了汇总。由试验结果可知,整个寿命周期内,包括低温初期工况和高温末期工况,镜头7温度在16~20.3 ℃之间,其余镜头温度在18.8~21.5 ℃之间,电子设备3~6温度在8.4~12.8 ℃之间,其余电子设备温度在2.6~27.3 ℃之间,音圈电机和步进电机温度在14~36 ℃之间,均满足温度指标要求。由试验结果还可看出,整个寿命周期内,镜头温度稳定度较高。同一工况下,镜头最大温度波动都在1 ℃以内,高于指标要求的4 ℃。从低温工况到高温工况,除位于入光口处的镜7外,其余镜头最大温度波动都在2.1 ℃以内,也高于指标要求的4 ℃。镜头7位于入光口,受冷黑空间、轨道外热流以及遮光罩温度影响较大,且为给音圈电机散热,在其附近布置有辐射冷屏,导致其低温工况到高温工况温度波动相对较大,为4.3 ℃,但也在其工作指标允许的6 ℃以内。监测仪光学镜头组件温度较高的稳定度,主要归功于两点,一是采用了间接热控思路,二则是通过结构热控一体化设计,在结构上充分保证了内热源系统与光学镜头室温之间的最大程度热隔离,最大程度削弱了内热源系统寿命周期内的较大温度波动(20 ℃,见表5)对光学镜头室温环境的影响。
图 10 高、低温初期瞬态工况监测仪主要部组件温度曲线
Figure 10. Temperature curves of main components on gas monitoring sensor under hot and cold case condition
表 5 试验结果汇总(单位: ℃)
Table 5. Summary of test results(Unit: ℃)
Components Cold
caseHot
caseTemperature
requirementOptical lens 1 18.8−19 20.4−20.9 20±2 Optical lens 2 20.8−21.0 20.8−21.0 20±2 Optical lens 3 19.6−19.8 19.8−20.0 20±2 Optical lens 4 21.3−21.5 21.0−21.3 20±2 Optical lens 5 19.9−20.0 19.9−20.1 20±2 Optical lens 6 20.2−20.5 20.5−21.1 20±2 Optical lens 8 19.1−20.0 20.4−21.2 20±2 Optical lens 7 16−16.6 19.1−20.3 Operating time19±3
Non-operating time19±7Electric device 1 9.2−12 14.2−19.1 −10−45 Electric device 2 19.8−20.4 26.7−27.3 −10−45 Electric device 3 8.4−9.7 9.5−12.4 0−20 Electric device 4 8.8−10.0 9.9−12.8 0−20 Electric device 5 8.7−9.9 9.9−12.7 0−20 Electric device 6 8.4−10.0 9.5−12.8 0−20 Electric device 7 2.6−4.5 17.1−21.7 −10−45 Electric device 8 4.8−6.5 20.1−24.2 −10−45 Electric device 9 4.5−6.6 20.0−24.1 −10−45 Electric device 10 11.1−12.5 22.3−26.0 −10−45 Electric device 11 9.1−10.8 19.0−24.8 −10−45 Voice coil motor 14.4−16.3 17.9−35.6 0−85 Stepper motor 18.5−18.6 28.7−35.9 −10−80
Thermal design of one space gas monitoring sensor and test validation
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摘要: 某空间气体监测仪结构布局紧凑,在较小尺寸空间内交错布置有8个镜头组件、11台电子设备内热源和2个电机。内热源数量众多,工作时间长,与镜头控温要求差别大,且1个电机为二维转动热源,这些特点给热设计带来挑战。为有效解决热控难题,采用了多种设计思路组合。基于热管理思路对监测仪各部组件热行为进行系统管理,以节省热控资源;基于间接热控思路对所处热环境复杂的光学镜头组件进行控温,提高其控温精度和温度稳定度;对转动电机则进行辐射冷却,避免在传热路径中引入挠性转动环节,以提高热控系统可靠性;并基于结构热控一体化设计,在结构上充分保证热设计各项需求。热平衡试验结果表明:高低温工况下,监测仪各部组件温度均满足指标要求,且整个寿命周期内,光学镜头温度稳定度较高,同一工况下光学镜头最大温度波动在1 ℃以内,实现了多热源复杂工作机制下光学镜头的高精度精密热控。Abstract: The structure layout of one space gas monitoring sensor is very compact. There are eight optical lens, eleven electronic devices and two motors staggered in the small-scale space. There were so many calorific equipments with long working hours and large power consumption, and their temperature control requirements were not consistent with the optical lens, the number of which was also large. Furthermore, one of the two motors was a two DOF turn motor during operating. These above characteristics make thermal design of the gas monitoring sensor a great challenge. To effectively solve the difficult problems of the gas monitoring sensor thermal design, combination of multy design methods were adopted. The thermal behavior of the gas monitoring sensor components were systematically managed based on the idea of thermal management to save thermal control resources. Indirect thermal control technology was used on the optical lens temperature control to guarantee meeting the high precision and stability requirement. Heat dissipation of the two DOF turn motor was achieved by radiation cooling, by which flexible rotating table could be avoided in the cooling path, so that thermal control system reliability could be improved. Finally, structural and thermal integrated design was applied to make sure the requirements of those above thermal design fully guaranteed in structure. The results of thermal balance test show that all components temperature meet the requirements no matter under cold case condition or hot case condition, and optical lens have high temperature stability throughout the life cycle. The maximum temperature fluctuation of all optical lens is less than 1 ℃ under the same case condition. High precision thermal control of optical lens are obtained under the condition of multiple heat source and complex working mechanism.
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表 1 内热源热耗及工作时间
Table 1. Heat consumption and operating time of calorigenic equipments
Calorigenic
equipmentsStand by
mode/WObservation
mode/WCalibration
mode/WOperating
time on
orbit/minElectric device 1 11 11 11 102 Electric device 2 0 3 3 68 Electric device 3 0 0.3 0.3 63 Electric device 4 0.9 1.2 1.2 102 Electric device 5 0.9 1.2 1.2 102 Electric device 6 0.9 1.2 1.2 102 Electric device 7 0 0 4 3 Electric device 8 0 20 20 63 Electric device 9 0 0 21.1 3 Electric device 10 8 8 8 102 Electric device 11 0 4 4 52 Voice coil motor 0 2 2 52 Stepper motor 0 1.7 1.7 52 表 2 气体监测仪各部组件控温要求
Table 2. Temperature control requirements of gas monitoring sensor components
Components Temperature requirement/℃ Operating time Non-operating time Optical lens 7 19±3 19±7 Optical lens 1–6,8 20±2 Voice coil motor of optical lens 7 0–85 Stepper motor of optical lens 7 –10–80 Electric device 1,2,7–11 –10–45 Electric device 3–6 0–20 表 3 气体监测仪在轨热边界条件及试验模拟方法
Table 3. Thermal boundary conditions of gas monitoring sensor on orbit and simulation methods of these conditions during ground test
Thermal boundary conditions Simulation methods during ground test Vacuum and space cold black background Use space environment simulator Thermal environment on satellite platform Design simulation tool of satellite platform,which is shown in Fig.9, Fig.5
Control the temperature of simulation tool according to the given temperature boundary of satellite;
Cover simulation tool surfaces facing gas monitoring sensor with MLI and put heating circuit on surface of the MLI to obtain orbit heat flux absorbed by MLIOrbit heat flux Use infrared heating cage and flux sensor to abtain heat flux incidenting to the entrance of earth baffle;
Put heating circuit on MLI surfaces and radiator back surface to obtain orbit heat flux absorbed by themHeat consumption of calorigenic equipments Calorigenic equipments of gas monitoring sensor operate during test according to the normal working mode on orbit 表 4 试验工况
Table 4. Operating conditions of test
Case Orbit heat flux One orbit working mode Temperature boundary
of satellite/℃Cold case Minimum heat flux throughout the life cycle Standby mode −5 Hot case Maximum heat flux throughout the life cycle Standby mode→calibration mode→
observation mode→standby mode45 表 5 试验结果汇总(单位: ℃)
Table 5. Summary of test results(Unit: ℃)
Components Cold
caseHot
caseTemperature
requirementOptical lens 1 18.8−19 20.4−20.9 20±2 Optical lens 2 20.8−21.0 20.8−21.0 20±2 Optical lens 3 19.6−19.8 19.8−20.0 20±2 Optical lens 4 21.3−21.5 21.0−21.3 20±2 Optical lens 5 19.9−20.0 19.9−20.1 20±2 Optical lens 6 20.2−20.5 20.5−21.1 20±2 Optical lens 8 19.1−20.0 20.4−21.2 20±2 Optical lens 7 16−16.6 19.1−20.3 Operating time19±3
Non-operating time19±7Electric device 1 9.2−12 14.2−19.1 −10−45 Electric device 2 19.8−20.4 26.7−27.3 −10−45 Electric device 3 8.4−9.7 9.5−12.4 0−20 Electric device 4 8.8−10.0 9.9−12.8 0−20 Electric device 5 8.7−9.9 9.9−12.7 0−20 Electric device 6 8.4−10.0 9.5−12.8 0−20 Electric device 7 2.6−4.5 17.1−21.7 −10−45 Electric device 8 4.8−6.5 20.1−24.2 −10−45 Electric device 9 4.5−6.6 20.0−24.1 −10−45 Electric device 10 11.1−12.5 22.3−26.0 −10−45 Electric device 11 9.1−10.8 19.0−24.8 −10−45 Voice coil motor 14.4−16.3 17.9−35.6 0−85 Stepper motor 18.5−18.6 28.7−35.9 −10−80 -
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