**4. Hybrid motor design & propellant performance**

This section presents the performance results for *C*32*H*66*=Mg=N*2*O=CO*<sup>2</sup> and *C*32*H*66*=Al=N*2*O=CO*<sup>2</sup> propellant combination. Theoretical specific impulse with respect to *O=F* (oxidizer to fuel) ratio is determined by NASA's Chemical Equilibrium Analysis (CEA) software [18]. Combustion energies of selected propellants are also explained. In addition, hybrid motor design is presented in this section. CAD drawing of the actual motor is presented accompanied with the real ignition process.

Carbon dioxide combustion slows down the chemical kinetics in the motor. This reduces the adiabatic flame temperature of the motor. The combustion boundary of experiments are depending on the adiabatic flame temperature limit. Therefore, the major goal of this section is to present the flame temperature change due to carbon dioxide addition in oxidizer mixture.

#### **4.1 Thermochemical analysis of propellant combination**

The **Figure 5** shows the *O=F vs Isp* for magnesium based propellants. Magnesium is loaded as 60% by mass. *CO*<sup>2</sup> mass fraction is 70% by mass in the oxidizer mixture. Chamber pressure, combustion efficiency, and area ratio are taken as 38 *bars*, 0*:*98 *and* 70. Ambient pressure is selected as 0.006 bar which is Martian atmospheric pressure value.

Magnesium and carbon dioxide addition to propellant combination clearly reduces the O *=F* ratio. Carbon dioxide increase in nitrous to 70% shifts to O*=F* ratio throughout 1. Reduction in O*=F* has the advantage of lower oxidizer mass tank. This further reduces the required nitrous mass brought from the Earth in a possible Mars Ascent Vehicle design.

The **Figure 6** shows the *O=F vs Isp* for aluminum based propellants. Al mass fraction is 40% in the paraffin. *CO*<sup>2</sup> mass fraction is 50% in the oxidizer mixture. Chamber pressure, combustion efficiency, and area ratio are taken as 38 *bars*, 0*:*90 *and* 70.

In **Figure 6**, although theoretical calculation assumes 90% combustion efficiency, practical *Al=CO*<sup>2</sup> based experiments show actual combustion efficiency of *Hybrid Propulsion System: Novel Propellant Design for Mars Ascent Vehicles DOI: http://dx.doi.org/10.5772/intechopen.96686*

**Figure 5.** *O=F ratio versus specific impulse – Magnesium based.*

#### **Figure 6.** *O=F ratio versus specific impulse – Aluminum based.*

70%. Therefore, theoretical *Isp* of aluminum-based experiments is found as 125 seconds at *O=F* ratio of 0.5. This shows that magnesium provides better performance and efficiency for MAVs.

### *4.1.1 Released energies of propellants*

Energy analysis is useful indicator of the propellant performance. The energy release due to the combustion is calculated by using heat of reactions ð*QR*Þ of chemical compounds (products and reactants). The Eq. 3 is the formulation for the heat of reaction (combustion) in kJ/kg for the specific propellant. In this equation, *Q*~ *<sup>c</sup>* is the heat of formations of products minus reactants. *nm* is the mole number and *MW* is the molecular weight of propellants (reactants).

$$Q\_R = \frac{\ddot{Q}\_c}{\sum \left( n\_m MW \right)\_{propellant}} \tag{3}$$

*Q*~ *<sup>c</sup>* refers the difference between total heat of formations of products and reactants at 25°C.

$$\tilde{\mathbf{Q}}\_{\mathbf{c}} = \sum\_{\text{products}} \Delta \tilde{H}\_{f \text{@25°C}} - \sum\_{\text{reactants}} \Delta \tilde{H}\_{f \text{@25°C}} \tag{4}$$


**Table 2.**

*Released Energy Values of Propellants.*

Energy analysis of various propellants presented in **Table 2**. Energy releases per kg of propellants is presented due to adiabatic flame temperature at the optimum O/F ratio. **Table 2** shows the fuel additive and carbon dioxide mass fractions and released energy ð*QR* ¼ *Eprop*Þ per kg of propellant. Results are negative that means reactions are exothermic.

**Table 2** indicates that aluminum provides 35 percent higher energy than the magnesium in *N*2*O* ignition. Furthermore, released energy of 70 % *CO*<sup>2</sup> based propellant is almost same as *Paraffin/Nitrous* propellant. This means that, although carbon dioxide slows down the chemical kinetics, it provides same energy level as *Paraffin*/*Nitrous*. Al and Mg provides similar heat of reaction for 70% CO2. Al and Mg with carbon dioxide provides higher heat of reaction than 70% CO2 based propellants.

#### **4.2 Rocket motor design**

A classical hybrid rocket motor is designed for the experiments. **Figure 7** illustrates the motor layout. The hybrid rocket motor consists of stainless-steel precombustion chamber with 3 grams solid fuel-based pyro section. The pyro is powered with 24 V battery. Pyro releases 5 kW to heat up the motor in 5 seconds. Brass injector chokes the flow therefore downstream pressure has no effect in flow rate. Brass retainer plate is significant to tether the injector. Oxidizer flow rate changes between 40 and 250 grams per second. Graphite insulator in precombustion chamber incarcerates the heat of the pyro fuel. There is a pressure transducer to measure downstream pressure (motor chamber pressure).

Combustion chamber has the 2.5 mm thick phenolic layer for the grain insulation. Single port fuel grain is 180 mm in length. Inner port diameter is 24 mm.

**Figure 7.** *Mars hybrid rocket motor lay-out.*

*Hybrid Propulsion System: Novel Propellant Design for Mars Ascent Vehicles DOI: http://dx.doi.org/10.5772/intechopen.96686*

And the outer grain diameter is 48 mm. Post combustion section has phenolic based layer to absorb the ignition. Nozzle (post combustion) pressure is also measured by the pressure transducer. Motor nozzle is made out from graphite. The throat diameter from 5 to 11 mm to regulate the combustion pressure.
