根据空间光学遥感器的轨道特点和任务需求，通过仿真分析对其进行了热设计。考虑近地空间环境的特殊性，选择防原子氧布作为多层隔热材料的面膜。为减小遥感器框架上安装的星上设备对遥感器温度的影响，设计了大热阻安装结构并使用了聚酰亚胺隔热垫。根据离轴三反光学遥感器及星载一体化卫星的结构特点，划分了主动加热区域，分配了加热功耗。由于遥感器对地观测频率低、工作功耗小、工作时间短，CCD焦面组件不设置散热面。根据遥感器的轨道参数和姿态，确定了3个典型工况并对其进行了仿真分析和热平衡试验。结果显示，遥感器本体温度为(18±4) ℃、光学元件温度为(18±2) ℃、CCD温度≤30 ℃，得到的仿真分析结果和试验数据验证了遥感器热设计的有效性。

A thermal simulation was established according to sensor parameters and the mission requirements to accomplish the thermal design of a lightweight space remote sensor. Atomic oxygen resistant cloth was chosen as the outmost material to reduce the damage by the space environment approximating to the earth. For some satellite devices installed on the back frame of the remote sensor, the connections with high heat resistance were designed and heat insulators were used to eliminate the heat influence on the devices. The positions and powers of the heaters were distributed according to the remote sensor′s structure characters. However, none radiator was set because of such small power and duty factor of the CCD components. Finally, the thermal design was certified by a thermal balance test. Three cases designed according to the orbit parameters and attitudes of the remote sensor were simulated and tested. The experiments show that the temperatures of frames and mirrors are (18±4) ℃ and (18±2) ℃, respectively, and the temperatures of the CCD components are lower than 30 ℃. The simulation analysis and the thermal balance test results both indicate that the thermal design is valid and content to the mission requirements.