随机近似热模型修正方法及相变热控关键问题研究
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摘要
热控系统是保证航天器正常工作的重要组成部分。航天器未来任务的发展推动了热控技术的进步,对热控系统的设计提出了新的要求。航天器热分析计算是实现航天器热控设计的必要手段。由于航天器内部换热的复杂性以及热分析模型某些输入参数的不确定性,往往需要对航天器热模型进行修正,以使热分析计算能够较准确地对航天器的运行特性和热设计水平进行评估。目前传统的热模型修正方法所面临的主要问题在于：过分依赖修正者的主观经验,修正效率与修正精度越来越难以满足航天器任务发展的需求,以及难以与现有热分析软件同步。因此,研究新的热模型修正方法对提高航天器的热设计水平具有重要意义。同时,作为航天器热控技术之一的相变热控技术虽然具有没有能耗、储能密度大、相变时温度近似恒定和经济性高等优点,但也存在热控相变材料封装困难、易发生泄露和热导率低等问题,因此,为了更好地利用相变材料实现热控,有必要对新型热控相变材料及其热性能进行研究。
本文工作重点在于从理论和实践方面研究新的热模型修正和热故障评估方法,解决现有热控相变材料所面临问题,探讨新型热控相变材料的实施方案。
首先,本文建立了采用蒙特卡洛随机近似方法结合优化算法的航天器热模型修正方法,并提出采用分层修正来提高热模型的修正精度。研究发现单纯的随机近似方法或局部优化算法都不能满足航天器热模型修正的需要,故本文建立了基于蒙特卡洛随机近似方法并结合优化算法的混合算法来进行航天器热模型修正；进一步研究发现,即使采用混合算法,也不能保证所有参数的修正精度,有些参数误差比较大,因此,本文提出通过随机近似多参数敏感性分析的方法对不确定参数对热控的影响程度进行定量分析,发现了多参数同时修正不能保证所有参数修正精度的原因。在此基础上,提出采用参数分层修正来提高修正精度的方法,即先修正对设备温度影响较大、与设备温度相关性较高的全局关键参数,在此基础上再进一步修正局部关键参数,然后再对其它参数进行修正。针对一虚拟卫星的热模型修正结果表明,采用分层修正时,热模型修正的精度远高于常规修正方法。进而针对某一地面试验状态下的模拟热控星,采用稳态工况1的试验数据进行热模型修正,并利用瞬态工况和稳态工况2的试验结果来验证修正后的热模型的准确性：修正后的热模型计算温度与试验温度较为吻合,计算温度与试验温度的偏差全部在±3。C以内。分析表明,本文所建立的航天器热模型修正方法克服了传统修正方法完全依赖设计经验的缺点,提高了热模型修正的精度和效率,可以很好地满足航天器热模型修正的要求。
在上述热模型修正方法的基础上,建立了航天器热故障参数评估方法,并对一地面试验模拟热控星的热故障参数进行了评估分析。利用该方法对受到破坏后的航天器热参数进行评估,得出受到不同程度破坏后的表面涂层热光学性质以及材料热物性等参数的分析值和变化规律；将各个工况温度的分析结果与试验结果以及热参数的评估值与试验值进行了对比,以此来验证热参数评估结果的有效性。分析结果表明,通过热故障参数评估方法得到的热参数分析结果能够很好地反映模拟热控星被破坏后的热参数的变化,涂层表面性质计算值与试验测量值吻合良好,最大相对误差小于15%,修正后的热模型计算得到的温度结果与试验结果误差较小,计算温度与试验温度的最大偏差为0.5℃,标准偏差最大值为1.9℃,95%以上的设备和舱板测点的温度差绝对值小于3℃。最后,根据热参数计算结果对模拟热控星在飞行状态下的热故障效应进行评估研究。
为了尽量避免各种不可预见的破坏对航天器热控的影响,有必要找出极端的热故障工况。由于完全通过试验方法找出极端热故障工况面临着较大的困难,因此本文将遭到破坏后模拟热控星的热参数变化情况对两颗典型卫星展开外推分析,研究了被破坏部位、破坏强度等因素对航天器热控的影响,分析发现,并不是破坏强度越高,热故障效果越好：非散热面被低强度破坏时整星设备温度最高,辅助散热面被中强度破坏时的温度升高水平与高强度破坏时相当；非散热面与辅助散热面组合时的热故障效应较好；选择极端故障工况的关键是对单个面被破坏后的热故障效果进行评估。
为了解决传统热控相变材料储能所面临的封装困难、传热性能差等问题,本文提出采用无需封装的高导热定形相变材料实现热控的方案。制备出以正十八烷为相变材料、以高密度聚乙烯为支撑材料、膨胀石墨作为导热增强剂的高导热定形相变材料,并对其热物性、相变特性、热稳定性以及相变储能特性进行了实验测量和理论分析,研究发现：该材料热导率为1.76W.m-1.K-1,比传统定形相变材料提高了4倍多,比纯石蜡提高了近11倍；材料的相变特性与正十八烷相差不大,该材料热性能稳定,真空质损率较低。数值模拟研究发现,该材料具有良好的储/放热性能,能够满足电子设备的热控制要求。
本文最后提出了利用高导热定形相变材料来防护短期高密度热流的方案。针对一颗虚拟卫星,分析了在常规及短期高热流条件下使用定形相变材料来改善航天器在高热流环境下抗冲击性的效果,并和常规的热控系统进行了对比,从理论上验证了这种方案的可行性。此外还研究了热导率对相变材料使用效果的影响。
Thermal control system is the system to guarantee the normal work of spacecraft. The development of spacecraft future mission promotes the progress of the thermal control technology advancement, and also puts forward new requirements to the design of the thermal control system. The spacecraft thermal analysis calculation is the necessary means for the realization of spacecraft thermal control design. Because of the complexity of heat transfer in spacecraft and uncertainty of some input parameters in thermal analysis model, the spacecraft thermal model often requires correction in order to make accurate evaluation of thermal design level and work characteristics of spacecraft. The traditional thermal model correction method is confronted with the following problems:1) It excessively depends on subjective experience of researchers;2) Its correction efficiency and precision is difficult to meet the needs of the development of the spacecraft mission;3) It is hard to synchronize with thermal analysis software. Therefore, developing new thermal model correction methods has great significance for the design level of spacecraft. As a thermal control technology of spacecraft, the phase change thermal control is featured by no energy consumption, high energy storage density, approximate constant temperature during phase change, high economy and so on. But it also has problems of needing tight packaged to prevent leakage and lower thermal conductivity. Therefore, in order to make better use of phase change materials for thermal control, it is necessary to develop new thermal control phase change materials and study its thermal performances.
This paper focuses on the study of new thermal model correction and thermal fault evaluation method from the theory and the practice angles, solving the problems of the traditional thermal control phase change materials, and discussing the thermal control design with new phase change material.
Firstly, the spacecraft thermal model correction methods are built using Monte Carlo stochastic approximation method combined with optimization algorithms, and layered correction method is proposed to improve the correction accuracy of the thermal model. The study has found that stochastic approximation algorithm and local optimization algorithm can not satisfy the requirements for thermal model correction when the uncertain parameters are corrected at the same time. Further study shows that hybrid algorithm also can not guarantee correction accuracy of all the parameters, and some parameters have large error. Therefore, the stochastic approximation multi-parameter sensitivity analysis is proposed to make quantitative analysis of the impact of the uncertain parameters on the thermal control, discovering the reasons why all parameters correction accuracy can not be guaranteed when they are corrected at the same time. In order to solve the problem, a layered correction method is proposed. In this method global key parameters are corrected firstly, and on this basis local key parameters are corrected. The results of a virtual satellite obtain by layered correction prove that the method in this study is superior to traditional methods. And thus a thermal model for a thermally controlled satellite in ground test conditions is corrected using above method. The test data for steady-state condition1are used to correct the thermal model, while the transient condition and steady-state condition2verify the corrected model. It is found that the calculated temperatures of the thermal model are identical with the test temperatures and the deviations between them are all within±3℃. The analysis show that the method in this study overcomes the disadvantage of traditional correction method, improves the precision and efficiency of thermal model correction, and overall it is successful in spacecraft thermal model correction.
On the basis of the above thermal model correction method, Thermal fault parameter evaluation method is built in the study. By the method, calculated values and variation rules of thermal-optical properties of surface coatings and thermal properties of materials are obtained while a thermally controlled satellite is damaged under different conditions. Calculation temperatures and the test results in different conditions are compared to verify the effectiveness of thermal fault parameters obtained. Also the thermal fault parameters obtained by calculation are compared to the experimental results. The results show that the calculated values obtained by the thermal fault parameter evaluation method are reasonable. And the results can accurately reflect the thermal properties changes when a satellite is damaged. The calculated values of the coating surface properties are in good agreement with the experimental values, and the max relative error is less than15%. The maximum error between the experimental values and the calculated values of temperature by the corrected thermal model is0.5℃, and the maximum standard deviation is1.9℃. The deviations between the calculated and test values of95%of the equipment and deck measuring points are all within±3℃. In the end, the thermal fault effect of thermally controlled satellite in orbit operation is studied according to the calculated value of thermal fault parameters.
In order to avoid the unpredictable damages to the spacecraft's thermal control, it is necessary to find out the extreme thermal fault conditions. It is hard to find out the extreme thermal fault conditions by experimental methods. Therefore, this paper conducts extrapolate applications of thermal parameter of a thermally controlled satellite damaged in ground test conditions to two typical satellites, and the effect of the damaged parts and strength on the spacecraft's thermal control is studied. The result shows that the higher damaged strength can not always get better thermal fault effect than lower damaged strength. When the non-radiative areas suffer low strength damage, the spacecraft temperature is higher. When the auxiliary radiative areas suffer moderate strength damage, the spacecraft temperature is almost identical to that suffering high damaged strength. The spacecraft has good thermal fault effect when the non-radiative areas and the auxiliary radiative areas are combined to be damaged. In order to find out the extreme thermal fault conditions, it is important to evaluate the thermal fault when a single area is damaged.
In order to solve the problems of packaging difficulty and low thermal conductivity in traditional phase change materials, a shape-stabilized phase change material with high thermal conductivity is proposed to be used for phase change thermal control. Using organic n-alkanes as phase change material, high-density polyethylene (HDPE) as support material, and expanded graphite (EG) as thermal conductivity enhancer, a new shape-stabilized phase change material with higher thermal conductivity is prepared. And its thermal property properties, phase change properties, thermal stability characteristics and heat storage/release performances are tested and analyzed. The research shows that the material has a thermal conductivity of1.76W.m-1.K-1, which increases over4times than that of the traditional shape-stabilized phase change material, and nearly increased11times than that of pure paraffin. There are no significant differences between the phase change properties of the material and pure paraffin. The material has good thermal stability and low mass loss percentage under vacuum condition. Also the material shows good heat storage/release performance, and is effective for electronic device thermal control.
In the last part of this paper, a solution of using a shape-stabilized phase change material with high thermal conductivity to protect the spacecraft suffering high energy is proposed. Taking a satellite as example, the thermal responses for spacecraft with shape-stabilized phase change material are investigated in contrast to that with conventional thermal control system, and it is confirmed that the solution given in the work is feasible. Also, the importance of thermal conductivity of phase change material on its application effect is discussed.
本文工作重点在于从理论和实践方面研究新的热模型修正和热故障评估方法,解决现有热控相变材料所面临问题,探讨新型热控相变材料的实施方案。
首先,本文建立了采用蒙特卡洛随机近似方法结合优化算法的航天器热模型修正方法,并提出采用分层修正来提高热模型的修正精度。研究发现单纯的随机近似方法或局部优化算法都不能满足航天器热模型修正的需要,故本文建立了基于蒙特卡洛随机近似方法并结合优化算法的混合算法来进行航天器热模型修正；进一步研究发现,即使采用混合算法,也不能保证所有参数的修正精度,有些参数误差比较大,因此,本文提出通过随机近似多参数敏感性分析的方法对不确定参数对热控的影响程度进行定量分析,发现了多参数同时修正不能保证所有参数修正精度的原因。在此基础上,提出采用参数分层修正来提高修正精度的方法,即先修正对设备温度影响较大、与设备温度相关性较高的全局关键参数,在此基础上再进一步修正局部关键参数,然后再对其它参数进行修正。针对一虚拟卫星的热模型修正结果表明,采用分层修正时,热模型修正的精度远高于常规修正方法。进而针对某一地面试验状态下的模拟热控星,采用稳态工况1的试验数据进行热模型修正,并利用瞬态工况和稳态工况2的试验结果来验证修正后的热模型的准确性：修正后的热模型计算温度与试验温度较为吻合,计算温度与试验温度的偏差全部在±3。C以内。分析表明,本文所建立的航天器热模型修正方法克服了传统修正方法完全依赖设计经验的缺点,提高了热模型修正的精度和效率,可以很好地满足航天器热模型修正的要求。
在上述热模型修正方法的基础上,建立了航天器热故障参数评估方法,并对一地面试验模拟热控星的热故障参数进行了评估分析。利用该方法对受到破坏后的航天器热参数进行评估,得出受到不同程度破坏后的表面涂层热光学性质以及材料热物性等参数的分析值和变化规律；将各个工况温度的分析结果与试验结果以及热参数的评估值与试验值进行了对比,以此来验证热参数评估结果的有效性。分析结果表明,通过热故障参数评估方法得到的热参数分析结果能够很好地反映模拟热控星被破坏后的热参数的变化,涂层表面性质计算值与试验测量值吻合良好,最大相对误差小于15%,修正后的热模型计算得到的温度结果与试验结果误差较小,计算温度与试验温度的最大偏差为0.5℃,标准偏差最大值为1.9℃,95%以上的设备和舱板测点的温度差绝对值小于3℃。最后,根据热参数计算结果对模拟热控星在飞行状态下的热故障效应进行评估研究。
为了尽量避免各种不可预见的破坏对航天器热控的影响,有必要找出极端的热故障工况。由于完全通过试验方法找出极端热故障工况面临着较大的困难,因此本文将遭到破坏后模拟热控星的热参数变化情况对两颗典型卫星展开外推分析,研究了被破坏部位、破坏强度等因素对航天器热控的影响,分析发现,并不是破坏强度越高,热故障效果越好：非散热面被低强度破坏时整星设备温度最高,辅助散热面被中强度破坏时的温度升高水平与高强度破坏时相当；非散热面与辅助散热面组合时的热故障效应较好；选择极端故障工况的关键是对单个面被破坏后的热故障效果进行评估。
为了解决传统热控相变材料储能所面临的封装困难、传热性能差等问题,本文提出采用无需封装的高导热定形相变材料实现热控的方案。制备出以正十八烷为相变材料、以高密度聚乙烯为支撑材料、膨胀石墨作为导热增强剂的高导热定形相变材料,并对其热物性、相变特性、热稳定性以及相变储能特性进行了实验测量和理论分析,研究发现：该材料热导率为1.76W.m-1.K-1,比传统定形相变材料提高了4倍多,比纯石蜡提高了近11倍；材料的相变特性与正十八烷相差不大,该材料热性能稳定,真空质损率较低。数值模拟研究发现,该材料具有良好的储/放热性能,能够满足电子设备的热控制要求。
本文最后提出了利用高导热定形相变材料来防护短期高密度热流的方案。针对一颗虚拟卫星,分析了在常规及短期高热流条件下使用定形相变材料来改善航天器在高热流环境下抗冲击性的效果,并和常规的热控系统进行了对比,从理论上验证了这种方案的可行性。此外还研究了热导率对相变材料使用效果的影响。
Thermal control system is the system to guarantee the normal work of spacecraft. The development of spacecraft future mission promotes the progress of the thermal control technology advancement, and also puts forward new requirements to the design of the thermal control system. The spacecraft thermal analysis calculation is the necessary means for the realization of spacecraft thermal control design. Because of the complexity of heat transfer in spacecraft and uncertainty of some input parameters in thermal analysis model, the spacecraft thermal model often requires correction in order to make accurate evaluation of thermal design level and work characteristics of spacecraft. The traditional thermal model correction method is confronted with the following problems:1) It excessively depends on subjective experience of researchers;2) Its correction efficiency and precision is difficult to meet the needs of the development of the spacecraft mission;3) It is hard to synchronize with thermal analysis software. Therefore, developing new thermal model correction methods has great significance for the design level of spacecraft. As a thermal control technology of spacecraft, the phase change thermal control is featured by no energy consumption, high energy storage density, approximate constant temperature during phase change, high economy and so on. But it also has problems of needing tight packaged to prevent leakage and lower thermal conductivity. Therefore, in order to make better use of phase change materials for thermal control, it is necessary to develop new thermal control phase change materials and study its thermal performances.
This paper focuses on the study of new thermal model correction and thermal fault evaluation method from the theory and the practice angles, solving the problems of the traditional thermal control phase change materials, and discussing the thermal control design with new phase change material.
Firstly, the spacecraft thermal model correction methods are built using Monte Carlo stochastic approximation method combined with optimization algorithms, and layered correction method is proposed to improve the correction accuracy of the thermal model. The study has found that stochastic approximation algorithm and local optimization algorithm can not satisfy the requirements for thermal model correction when the uncertain parameters are corrected at the same time. Further study shows that hybrid algorithm also can not guarantee correction accuracy of all the parameters, and some parameters have large error. Therefore, the stochastic approximation multi-parameter sensitivity analysis is proposed to make quantitative analysis of the impact of the uncertain parameters on the thermal control, discovering the reasons why all parameters correction accuracy can not be guaranteed when they are corrected at the same time. In order to solve the problem, a layered correction method is proposed. In this method global key parameters are corrected firstly, and on this basis local key parameters are corrected. The results of a virtual satellite obtain by layered correction prove that the method in this study is superior to traditional methods. And thus a thermal model for a thermally controlled satellite in ground test conditions is corrected using above method. The test data for steady-state condition1are used to correct the thermal model, while the transient condition and steady-state condition2verify the corrected model. It is found that the calculated temperatures of the thermal model are identical with the test temperatures and the deviations between them are all within±3℃. The analysis show that the method in this study overcomes the disadvantage of traditional correction method, improves the precision and efficiency of thermal model correction, and overall it is successful in spacecraft thermal model correction.
On the basis of the above thermal model correction method, Thermal fault parameter evaluation method is built in the study. By the method, calculated values and variation rules of thermal-optical properties of surface coatings and thermal properties of materials are obtained while a thermally controlled satellite is damaged under different conditions. Calculation temperatures and the test results in different conditions are compared to verify the effectiveness of thermal fault parameters obtained. Also the thermal fault parameters obtained by calculation are compared to the experimental results. The results show that the calculated values obtained by the thermal fault parameter evaluation method are reasonable. And the results can accurately reflect the thermal properties changes when a satellite is damaged. The calculated values of the coating surface properties are in good agreement with the experimental values, and the max relative error is less than15%. The maximum error between the experimental values and the calculated values of temperature by the corrected thermal model is0.5℃, and the maximum standard deviation is1.9℃. The deviations between the calculated and test values of95%of the equipment and deck measuring points are all within±3℃. In the end, the thermal fault effect of thermally controlled satellite in orbit operation is studied according to the calculated value of thermal fault parameters.
In order to avoid the unpredictable damages to the spacecraft's thermal control, it is necessary to find out the extreme thermal fault conditions. It is hard to find out the extreme thermal fault conditions by experimental methods. Therefore, this paper conducts extrapolate applications of thermal parameter of a thermally controlled satellite damaged in ground test conditions to two typical satellites, and the effect of the damaged parts and strength on the spacecraft's thermal control is studied. The result shows that the higher damaged strength can not always get better thermal fault effect than lower damaged strength. When the non-radiative areas suffer low strength damage, the spacecraft temperature is higher. When the auxiliary radiative areas suffer moderate strength damage, the spacecraft temperature is almost identical to that suffering high damaged strength. The spacecraft has good thermal fault effect when the non-radiative areas and the auxiliary radiative areas are combined to be damaged. In order to find out the extreme thermal fault conditions, it is important to evaluate the thermal fault when a single area is damaged.
In order to solve the problems of packaging difficulty and low thermal conductivity in traditional phase change materials, a shape-stabilized phase change material with high thermal conductivity is proposed to be used for phase change thermal control. Using organic n-alkanes as phase change material, high-density polyethylene (HDPE) as support material, and expanded graphite (EG) as thermal conductivity enhancer, a new shape-stabilized phase change material with higher thermal conductivity is prepared. And its thermal property properties, phase change properties, thermal stability characteristics and heat storage/release performances are tested and analyzed. The research shows that the material has a thermal conductivity of1.76W.m-1.K-1, which increases over4times than that of the traditional shape-stabilized phase change material, and nearly increased11times than that of pure paraffin. There are no significant differences between the phase change properties of the material and pure paraffin. The material has good thermal stability and low mass loss percentage under vacuum condition. Also the material shows good heat storage/release performance, and is effective for electronic device thermal control.
In the last part of this paper, a solution of using a shape-stabilized phase change material with high thermal conductivity to protect the spacecraft suffering high energy is proposed. Taking a satellite as example, the thermal responses for spacecraft with shape-stabilized phase change material are investigated in contrast to that with conventional thermal control system, and it is confirmed that the solution given in the work is feasible. Also, the importance of thermal conductivity of phase change material on its application effect is discussed.
引文
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[12]Cheng W L, Ma R, Xie K, et al. Simultaneous measurement of thermal properties by thermal probe using Stochastic approximation method. Applied Energy,2011,88(5):1834-1840
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[24]Khateeb S A, Amiruddin S, Farid M, et al. Thermal management of Li-ion battery with phase change material for electric scooters:Experimental validation. Journal of Power Sources,2005, 142(1-2):345-353
[25]Tan F L, Fok S C. Thermal Management of Mobile Phone using Phase Change Material. IEEE 9th Electronics Packaging Technology Conference, Dec 10-12, Singapore (2007), pp.836-842
[26]Fok S C, Shen W, Tan F L. Cooling of portable hand-held electronic devices using phase change materials in finned heat sinks. International Journal of Thermal Sciences,2010,49(1):109-117
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