**6. Hybrid propulsion for Mars ascent vehicles**

Mars Ascent Vehicle design concepts are studied by many researchers. Carter [21] discussed technology requirements for propulsion systems such as displacement pumps and bladder lined composite tanks. One of the experimental works for a potential MAV is presented by Karp with paraffin-based fuel and Mixed Oxides of Nitrogen (MON-3) oxidizer [22]. Furthermore, Evans and Karabeyoglu also studied MON based oxidizer for MAV experiment with metallized SP7 fuel by using 30 microns sized aluminum powder [23]. SP7 paraffin-based solid fuel is developed by Space Propulsion Group, Inc., specifically for this program.

This chapter proposes a classical hybrid propulsion system by using *Paraffin=Mg=CO*2*=N*2*O* propellant combination is totally feasible for MAVs. Propellant combination provides significant cost savings as well as ease of manufacturing. Self-pressurizing capability of oxidizers also make the system simple and safe. Reduction in oxidizer to fuel ratio is another advantage of the system that reveals lighter oxidizer tank.

A hybrid rocket motor with a 38 bars combustion pressure, 98% of combustion efficiency, 3 *kg=s* of oxidizer mass flow rate and nozzle area ratio of 70 provides 250 seconds of specific impulse at an oxidizer to fuel ratio ð Þ *O=F* of 1.1. Also, the thrust level of this single motor is found as 13 kN. Therefore, a potential Mars Sounding Rocket with total fuel mass of 40 kg have total oxidizer mass of 44 kg. Total propellant mass corresponds 84 kg. 1 kg payload mass is assumed to transfer a sample from a point to another on the Mars. Structure mass of the rocket is estimated as 18 kg consist of motor interface, motor casing, measurement devices and the oxidizer feed system (plumbing and valves). Avionics and power system is 5 kg. Nosecone is 2 kg.

The oxidizer tank mass is depending on the density of the oxidizer at 50 bars. It is worth to note that saturated oxidizer mixture at 50 bars need the tank temperature of 15<sup>∘</sup> C. Mars atmospheric temperatures changing between �<sup>70</sup> *and* <sup>20</sup><sup>∘</sup> C near the equator. Thus, a heater system with a simple thermostat can be used at night times to increase the temperature to 15<sup>∘</sup> C. On the other hand, launch can be performed in Martian summer days such as July or August.

At the 50 bars, *CO*2*=N*2*O* mixture has a liquid phase density of 816 *kg=m*3. 44 kg of oxidizer makes 54 liters of oxidizer tank for the liquid phase. Therefore, tank mass with 4 mm thickness is 8.7 kg.

The sounding rocket has 17 seconds burn time and the mass ratio *<sup>n</sup>* <sup>¼</sup> *Minitial Mfinal* of 4.8. 2-DOF calculation shows that rocket have 47 km downrange distance with 24 km burn out altitude. The maximum altitude achieved is 2 km. However, 3-DOF calculation is needed in order to increase the launch precision. The proposed rocket system can be scaled up as Mars Ascent Vehicle. By using the same performance parameters, Δ*V* of MAV system is found as 3850 m/s. In addition, Δ*V* requirement of 500 km low Martian orbit is found as 3652.3 m/s. It is worth to note that drag force is neglected since the calculation needs the drag coefficient. Drag coefficient will be analyzed in future study as the MAV trajectory design.

**Figure 14.** *Oxide formation after the motor experiment.*

Potential MAV design needs at least two staged rocket design to fulfill 3652 m/s delta-V requirement. Because specific impulse of propellant selection is quite low for a single staged rocket. Thus, single staged rockets can be used for Mars ballistic hopper missions (such as Mars Sounding Rocket). In addition, oxidizer mass which is needed to brought from the Earth should be minimized for practical in-situ Mars sounding rocket. Considering a single staged hopper rocket with 1600 m/s reveals following minimization process. Eq. (5) is used for this minimization process.

$$\frac{\mathbf{M}\_{earth}}{\mathbf{M}\_{pl}} = \left[\frac{e^{\Delta V\_{dd}/l \text{pg}\_0} - \mathbf{1}}{\mathbf{1} - \epsilon e^{\Delta V\_{dd}/l \text{pg}\_0}}\right] \left[\epsilon + \left[\frac{\mathbf{1} - \epsilon}{\mathbf{1} + O/F} (\mathbf{1} + a\mathcal{O}/F)\right]\right] \tag{5}$$

*ϵ* is the structural mass fraction. *α* is the nitrous oxide mass fraction in the oxidizer mixture. Nitrous mass fraction is the key parameter refers the earth based mass. O/F ratio is taken from CEA.

Earth based mass is minimized due to the payload mass. **Figure 14** summarizes minimized *Mearth*/*Mpl* due to nitrous oxide mass fraction for magnesium powder cases.

Although practical experiments are performed due to the Mg60 based fuel grains, Mg80 shows minimum mass fraction. Minimum values are taken at optimum O/F ratios. In addition, **Figure 14** shows minimum values both for Al80 and Mg80. Aluminum provides 5% smaller fraction than the magnesium. However, Al80 ignition with *CO*2. is not practical.

### **7. Conclusion**

This book chapter aims the explain the fundamentals of the hybrid propulsion system. In addition, readers can also be understand the practical setup configuration in hybrid rockets. Experimental tests are performed in order to understand the combustion characteristics of the carbon dioxide. Paraffin wax based fuel is the main binder. Aluminum and magnesium are selected as fuel additive. *CO*2*=N*2*O* mixture provides sustainable combustion mechanism.

This work concludes that *Paraffin=Mg=CO*2*=N*2*O* is the most feasible propellant for Martian rockets. Carbon dioxide addition to the propellant reduces the oxidizer to fuel ratio. This means that Mars rockets can use lighter oxidizer tank. In-situ *CO*<sup>2</sup> *Hybrid Propulsion System: Novel Propellant Design for Mars Ascent Vehicles DOI: http://dx.doi.org/10.5772/intechopen.96686*

significantly reduces the mass needed to brought from the Earth. Self-pressurizing feature of the *CO*2*=N*2*O* reduces complexity and the cost of the rocket. In addition, *Mg=CO*<sup>2</sup> combustion provides really high motor efficiencies. Hybrid rockets that uses in-situ based *Paraffin=Mg=CO*2*=N*2*O* is the most prominent candidate due to several aspects such as low cost, safe launch operations, ease of manufacturing, and ease of design.
