# Aerospace
## 8 Mission Cost  
[PERSON] et al. performed a preliminary study assessing the cost drivers for small missions and concluded that \(\Delta V\) is the largest driver of mission cost [7]. Figure 32 shows the estimated mission cost as a function of \(\Delta V\) for a chemical propulsion system and payload mass. The 25 kg spacecraft design considered in this study could potentially carry about 5 kg of scientific payload (based on the study by [PERSON] et al. [42]).  
Compared to the 1770 m/s required for propulsive insertion to a 200 \(\times\) 2000 km orbit at Mars, aerocapture accomplishes orbit insertion with less than 100 m/s (combining PRM and ACM \(\Delta V\); cf. Section 4.7). This translates to a cost reduction from over USD 100 million to USD 20 million, nearly a factor of five. At Venus, compared to the 3130 m/s required for propulsive insertion, aerocapture requires less than 60 m/s (cf. Section 5.7), thus potentially providing even more cost savings. A caveat of this estimate is that the cost of the aerocapture system is not considered, as no such estimate is available in the literature. The realization of a drag-modulation aerocapture flight system will involve  
Figure 31: Venus small satellite constellation in 200 \(\times\) 2000 km near-polar inclination orbits (green), and the host orbiter in a 250 \(\times\) 150,000 km polar orbit (magenta).  
Figure 32: Mission cost as a function of \(\Delta V\) for a chemical propulsion system. \(m_{\rm PL}\) refers to the science payload (instrument mass). Derived using data from [PERSON] et al. [7].  
hardware and software development and testing, validation, and integration costs, which cannot be estimated at this stage. Once a viable design is realized and flight tested, the recurring cost of realizing additional flight units and integrating them with small satellites is expected to be much lower than the initial development costs. Hence, the significant \(\Delta V\) saving offered by aerocapture will potentially lead to cost savings for small missions, and additional studies are recommended to quantify the cost for small satellite aerocapture mission architectures.  
Another method of comparing aerocapture to propulsive insertion is to compare the orbit insertion fuel mass to the aerocapture system mass, as the required fuel and aerocapture system mass can be an indicator of the cost. For propulsive insertion, using the rocket equation, given a spacecraft dry mass \(m_{d}\), the required fuel mass \(m_{f}\) is  
\[m_{f}=m_{d}\bigg{(}\exp\bigg{[}\frac{\Delta V}{I_{sp}g_{0}}\bigg{]}-1\bigg{)} \tag{10}\]  
For insertion to a 200 \(\times\) 2000 km orbit at Mars, the \(\Delta V\) required is 1770 m/s. Assuming \(I_{sp}\) = 320 s for a bi-propellant system, an impulsive burn and no finite burn losses, the required \(m_{f}\) to insert a 25 kg dry mass spacecraft into orbit is 19 kg. For a 200 \(\times\) 300 km orbit, the \(\Delta V\) is 2100 m/s, and the required fuel mass is 23 kg. The aerocapture system mass is approximately 25 kg, irrespective of the orbit size. Hence, the aerocapture does not lead to mass savings over propulsive capture for orbit insertion at Mars for typical Earth-Mars trajectories. Once a drag modulation flight system has been realized through technology development, integrating a 25 kg compactly stowable aerocapture system may be more cost-effective than a 25 kg propulsion system for a small spacecraft with volume constraints.  
For insertion to a 200 \(\times\) 2000 km orbit at Venus, the \(\Delta V\) required is 3130 m/s. The required \(m_{f}\) to insert a 25 kg spacecraft into orbit is 43 kg, which is 72% higher than the aerocapture system, which is 25 kg. For a 200 \(\times\) 300 km orbit, the \(\Delta V\) is 3500 m/s and the required fuel mass is 51 kg, which is over 100% more than the aerocapture system mass. Hence, the aerocapture offers significant mass savings over the propulsive capture at Venus, particularly to low-circular orbits, which will potentially also lead to cost savings.