Study On Stochastic Approximation Thermal Model Correction Method And Phase Change Thermal Control

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.