## 2 The Impact of Satellite Bus Thermal Deformation on the Laser Communication Load Acquisition Time
The distribution of this part of the error is listed in Table 2. The uncertainty error is calculated by the conservative method of absolute sum, and the formula is shown in (2). \[\Delta=\sum_{i=1}^{m}|e_{i}| \tag{2}\] \begin{table}
\begin{tabular}{c c c} \hline
**ID** & **Parameters** & **Specifications** \\ \hline
1 & Mode & Full-duplex \\
2 & Transmission rate @ 5000 km & 1.25 Gbps \\
3 & Transmission rate @ 3000 km & 2.5 Gbps \\
4 & Sensitivity & \(\leq\)\(-\)40 [EMAIL_ADDRESS] Gbps, \(\leq\)\(-\)37 [EMAIL_ADDRESS] Gbps \\
5 & Error rate & \(\leq\)\(1\times 10^{-7}\) (After decoding) \\
6 & Tracking range & 200–4500 km \\
7 & Precise tracking accuracy & \(\leq\)5 μrad \\ \hline \end{tabular}
\end{table} 
Table 1: Parameters of the laser communication load. The satellite attitude measurement error in Table 2 is affected by the star sensor attitude error, the star sensor temperature drift error, the thermal deformation error of the satellite bus, and other factors. The thermal deformation error of the satellite bus plays a dominant role. When the thermal deformation error is controlled within 0.16 mrad, the attitude measurement error of the satellite can meet the index requirement of 0.35 mrad. It is challenging to regulate the thermal deformation error of the honeycomb sandwich communication satellite bus within 0.16 mrad due to a number of reasons. First off, because communication loads come in a variety of forms, operate in a variety of ways, and consume a lot of power, there is a significant disparity between the peak power and steady power of communication satellites. In order to lower the cost of launch, the thermal design of the satellite must take into account both the volume and weight of the entire satellite as well as an adequate cooling area to fulfill the peak cooling power requirement. The temperature variation of the satellite bus between normal power working mode and peak power working mode is caused by the difference between normal power and peak power as well as the difference in external heat flow while the satellite is at different points in orbit. Finding a steady satellite temperature distribution state to use as the temperature reference for the thermal deformation input is therefore challenging. The honeycomb sandwich structure's characteristics make it challenging for communication satellites to locate a stable thermal reference for the installation of high-precision attitude measurement components such as star sensors and laser communication loads. The aforementioned two points have the effect of altering the uncertain region where the laser communication load points by causing the angle variation between the laser communication load and the star sensor to fluctuate significantly in the classic design mode of laser communication satellites. In Section 4, it is discovered through ground tests that when the star sensor and laser communication load are installed conventionally on the satellite structural plate, the angle variation between the two equipment's prisms caused by the thermal deformation of the satellite bus will reach 117.74\({}^{\prime\prime}\) and the corresponding synthetic uncertainty will reach 0.57 mrad, much higher than the required 0.16 mrad. The capture time increased by 65.16% to 46.55 s, despite the uncertainty area expanding by 27.3%. This was due to a quadratic positive correlation between the capture time and the uncertainty area. The thermal deformation management of the satellite bus must be optimized in order to achieve the design requirement of laser communication load capture time within 30 s. Controlling the uncertain region of the laser communication load is crucial for quickly acquiring and building the link in the business mode of the laser communication satellite.