## 7 Conclusions  
The research successfully defined realistic propulsion, thermal and power systems for a 3U CubeSat carrying the RAVAN payload developed by NASA on board at an altitude of 600 km.  
The propulsion system was selected as a Mars-Space PPTCUP micro-pulsed plasma thruster. The suitability of this system was verified using first-principle hand calculations to ensure the thruster could provide enough impulse to overcome the generated force due to drag to maintain an altitude of 600 km and bring the CubeSat down to a graveyard orbit of 513 km. However, future work should be conducted into propellant storage and orbital maneuvers.  
The thermal system was analyzed using first-principle steady-state heat transfer calculations, and then compared to published literature and Analytical Computational analysis performed upon the SolidWorks thermal solver. A steady-state heat transfer case was developed for each methodology, generating results with negligible percentage error compared to those published in literature. The analysis found that during the eclipse phase, temperatures dropped as low as \(-80\)\({}^{\circ}\)C assuming only Earth's infra-red flux was incident upon the CubeSat and heat was radiating to background space at 3 K on all other 5 remaining faces. At such a low temperature, the CubeSat's electrical system would fail, causing mission failure. It was therefore suggested that the implementation of a heater was necessary to prevent temperatures exceeding the allowable operational range. The hot-case analysis found that the integration of a passive thermal system comprising of Black high-emissivity paint and MLI was required to prevent excessive heating within the structure.  
The power system analysis successfully defined electrical consumption scenarios for the CubeSat's 600 km orbit. The analysis verified that a singular 7 W solar panel mounted on a Sun-facing side of the CubeSat using a sun sensor could satisfactorily power the electrical system throughout the hot phase and charge the craft's battery enough to ensure constant electrical operation during the cold phase, even with the additional integration of an active thermal heater. However, when the inevitable end-of-life degradation of the solar cell was factored into the analysis it was found that singular panel could no longer power the electrical system throughout the entire orbit, leading to an approximate power deficit of 2 KW. To overcome this power deficit, it was suggested that a second solar cell be integrated into the antenna housing 100 mm \(\times\) 100 mm face. This will ensure the addition of another 2 W of theoretical power on top of the existing 7 W assuming both faces are operating simultaneously.  
Figure 59: Cold case power dissipation simulation.  
In future work, however, a detailed CAD model should be developed, ensuring all selected components can fit within the specified 3U volume. Further analysis should also be undertaken into the weight and position of center of gravity of any developed configuration. This detailed CAD model could then be integrated into the transient case solver in SolidWorks thermal analysis to determine the cyclic loading due to temperature change of the orbit cycle and pinpoint the exact required activation time of the CubeSat battery heater to optimize power consumption compared to running the heater for the whole cold eclipse cycle.  
[PERSON] performed the analysis and the simulation, [PERSON] and [PERSON] wrote the paper. The research was conducted under the supervision of [PERSON].  
This research received no external funding.  
The research was self-funded. The research was performed in the framework of [PERSON] engineering degree.  
The authors declare no conflict of interest.