# Effect of Structural Materials on Monopropellant Thruster Propulsion Performance in Micro Scale
## 4 Results and Discussion
### Heat Loss
A variation in  
Figure 5: Temperature at four different locations of the microthruster affected by variation in heat loss, depending on structural materials and propellant weight percentages (A: 95%, B: 90%, C: 85% H\({}_{2}\)O\({}_{2}\)). the sensitivity of the heat loss in relation to thermal diffusivity was also observed. In regard to this figure, the chamber heat loss proportional to the thermal diffusivity was established in the following manner:  
\[q_{c}\sim\left(\frac{k}{\rho c_{p}}\right)^{a} \tag{11}\]  
where \(a\) is 0.124, 0.126, and 0.130 for 85, 90, and 95 wt% hydrogen peroxide, respectively, in the case of thermal diffusivity below 3.7 mm\({}^{2}\)/s, and 0.0111, 0.0113, and 0.0118 for 85, 90, and 95% hydrogen peroxide, respectively, in the case of thermal diffusivity above 3.7 mm\({}^{2}\)/s. The chamber heat loss calculation based on the nozzle inlet temperature was followed by downstream heat loss calculation based on the nozzle throat and exit temperatures. With respect to nozzle throat heat loss, a similar form of equation was used, but for the nozzle throat, (T\({}_{\text{nozzle throat}}\) adiabatic \(-\) T\({}_{\text{nozzle throat}}\))/T\({}_{\text{nozzle throat}}\) adiabatic, which was the temperature difference at nozzle throat, for a range of thermal diffusivity, compared to the adiabatic case as a reference. Likewise, (T\({}_{\text{nozzle exit adiabatic}}\) \(-\) T\({}_{\text{nozzle exit}}\))/T\({}_{\text{nozzle exit adiabatic}}\) was used for nozzle exit thermal energy loss, as described in Figure 7 in the same scale for heat loss comparison between the measurement locations. Comparing the thermal losses in relation to the different propellant enthalpies and, thus, different inlet temperatures, it was found that the thermal energy loss measured at the nozzle inlet was larger than that at the nozzle throat, particularly in the low thermal diffusivity range, meaning there was more loss in the chamber than elsewhere, at least down to the nozzle throat of the thruster design considered. The difference between the nozzle inlet and throat, however, was more marginal as the thermal diffusivity increased, resulting in more thermal loss near the nozzle throat in regard to the high-thermal-diffusivity materials. This phenomenon mirrored the trend for heat transfer in rockets which have the highest heat flux at the nozzle throat and a much higher velocity than that of their chamber. The thermal loss measured at the nozzle exit was similar to that of the nozzle inlet at low thermal diffusivity, but at high thermal diffusivity, the loss was greatest at the nozzle exit across the nozzle, reaching around 15% in the case of 95 wt% H\({}_{2}\)O\({}_{2}\), meaning the thermal  
Figure 6: Heat transfer coefficient of the microthruster with materials of various levels of thermal diffusivity and different propellant enthalpies (85, 90, and 95 wt% hydrogen peroxide). loss at the nozzle exit was more sensitive to variations in thermal diffusivity than was the case in regard to the nozzle inlet and the chamber. Overall, the figure indicates a higher thermal energy loss with regard to the higher-enthalpy propellant, regardless of thermal diffusivity, and that the loss was greater between the propellants at higher thermal diffusivity and lower at lower thermal diffusivity.In other words, the thermal loss regarding the propellant enthalpies was more material-sensitive in the high thermal diffusivity range. It was also found that the thermal loss variation with respect to thermal diffusivity had different slopes, being more sensitive to thermal diffusivity variation in the low thermal diffusivity range. Figure 8 compares the area-averaged temperature on the outside surface of the solid with the chamber inlet temperature. In the thermal diffusivity range covered, the outside temperature ranged from 469.23 to 580.56 K for 85 wt% hydrogen peroxide, from 510.00 to 648.13 K for 90 wt%, and from 553.28 to 720.64 K for 95 wt%.