为解决有限热控资源下卫星多个光学遥感载荷及平台单机的热控问题, 对该卫星采取主、被动热控相结合的设计方案。首先, 根据卫星自身特点、热控需求及轨道外热流确定热设计的总体方案; 接着, 针对光学载荷和平台重要单机进行详细热设计说明, 并利用有限元分析软件计算卫星各组件的温度结果; 然后, 开展整星热平衡试验, 获取试验温度结果判断热设计的正确性; 最后, 通过对比卫星在轨遥测、热分析及热试验温度数据, 验证了该热设计方案的实际效果。在轨遥测数据显示: 主载荷相机温度控制在19.7~20.3 ℃之间, 光学小载荷温度控制在-31.2~6.6 ℃之间, 舱内单机温度在9.7~29.5 ℃之间。各温度结果均满足热控指标要求, 在轨数据与热分析及热试验结果偏差小于±3 ℃。表明该光学遥感卫星热设计正确可行, 热分析及热试验过程合理可靠。

In order to solve the thermal control problem of multiple optical remote sensing payloads and platform units on asatellite using limited thermal control resources, a design scheme based on active and passivethermal control strategies was presented. First, according to the satellite characteristics, thermal control requirements, and orbital heat flux, the thermal designing overall project was confirmed. Next, detailed thermal design instructions foroptical payloads and important platform instruments were listed, and the temperaturesof satellite subassemblies are calculated by finite element analysis software. Then, a thermal balance experiment on the whole satellite system was carried out to obtain test temperatures and verify the correctness of the thermal design. Finally, the real effect of the thermal design scheme was proved by comparing the temperature data obtained from on-orbit telemetry, thermal analysis, and thermal testing of the satellite. On-orbit telemetry data indicate that the temperature of the main payload camera is controlled from 19.7 ℃ to 20.3 ℃.The temperature of minor optical payloads ranged from -31.2 ℃ to 6.6 ℃,and the temperature of units inside the satellite cabinranged from 9.7 °C to 29.5 ℃. All the temperature results meet the requirements of the thermal control index. The temperature difference between on-orbit telemetry, thermal analysis, and thermal testing is less than ±3 ℃.The results show that the thermal design of the optical remote sensing satellite is correct and feasible, while the thermal analysis and test process are reasonable and credible.