## 5 Thermal Analysis and Results Discussion  
### Thermal Analysis of a 6U Satellite with Heat Pipes  
For the satellite configuration shown in Figures 14 and 15 and heat load conditions as listed in Table 7, thermal analyses were carried out for both worst hot and worst cold conditions with constant conductance heat pipes (CCHP). The thermal model of the 6U satellite with cryogenic payload is shown in Figure 16. This thermal model has a micro cryocooler modelled as a cylinder with a copper cold finger, generic cryogenic payload also modelled as a cylinder, and heat dissipating elements such as electronics, heaters, and batteries. The thermal contact conductance values used are provided in Table 6. For this study, it is considered that the satellite is orbiting around the earth at 535 km equatorial orbit with 5\({}^{\circ}\) inclination. As can be seen from the thermal model, one of the larger surfaces of the 6U satellite (maximum radiator area considered for BMR is 290 mm \(\times\) 250 mm) is used as the radiator plate to which the heat pipes are attached. This study considers a generic mid-wave infrared instrument whose detector must be kept around 95K as reference. To provide this cryogenic cooling, a miniature stirling cryocooler (Ricor K508) is considered as it provides better performance at low input power. Figure 16: 6U Satellite Body-mount radiator. \begin{table}
\begin{tabular}{c c c} \hline \hline
**Components** & **Peak Operation (W)—Hot** & **Nominal (W)—Cold** \\ \hline Electrical Power System (EPS) & 3.5 & 3.5 \\ ADCS & 0.65 & 0.65 \\ OBDH & 0.5 & 0.5 \\ Payload & 12.0 & 0.0 \\ Camera & 1.75 & 0.0 \\ Battery & 3.75 & 3.75 \\ cryocooler & 17 & 0 \\ \hline \hline \end{tabular}
\end{table} 
Table 7: Thermal Model—Heat loads. Figure 15: Satellite’s attitude in orbit—DR. 860 g. The temperature summary of the body-mount radiator with embedded heat pipes is shown in Figure 16. All the internal components are exchanging heat between themselves and the external plates of the satellite exchange the heat with external space environment. The amount of heat absorbed or emitted from any component is strongly dependent on its surface thermo-optical properties. The 6U satellite with BMR is analyzed for worst hot case orbital condition. The MWIR instrument in this study is assumed to take science measurement all the time in the orbit and hence the micro cryocooler considered to be running continuously to meet the instrument's thermal requirement at the detector. Cryocooler is mounted to the HX and heat pipes are connected to both HX and BMR. The main objective of this arrangement is to keep the temperature within the system interface temperature of 25 \({}^{\circ}\)C to 30 \({}^{\circ}\)C. However, the simulation results show that the body-mounted radiator panel with four heat pipes is not able to provide sufficient thermal performance as the temperature of the Cryocooler unit (CCU) exceeds its maximum operating temperature of 85 \({}^{\circ}\)C as seen in Figure 17. Since both CCU and HX are placed inside the satellite, their temperature should be maintained within the desirable limit in order to ensure the rest of the satellite components are safe for continuous satellite operation. An alternative thermal solution, deployable radiator, is studied for the same satellite configuration. 2. Deployable radiator: For the same 6U satellite thermal model, the radiator is changed to deployable configuration to compare the radiating capacity with BMR panel. As with the BMR, the deployable radiator is made from aluminum 6061-T6 with a thickness of 6 mm. CCHP are embedded to both HX and radiator separately and are linked together using thermal straps along the edges of the radiator panel to make it flexible. The surface area of the deployable radiator shown in Figure 18 is assumed to be 150 mm \(\times\) 270 mm with 620 g of mass including heat pipes.