**1. Introduction**

Mars is known to have the most suitable geological features and atmospheric conditions for the future human spaceflight. Based on the data from the orbiters and rovers sent through deep space, Mars has the most active volcanic mountains and the highest impact craters of all the worlds. Strong evidences via rover measurements moots that liquid water may have poured across the surface of Mars billions of years ago [1]. There is also evidence of methane leakage between rocks that could be an indication of microbial life. Furthermore, the location of Mars smooths the way of long-term human spaceflight. Mars is relatively close to Earth compared to other possible planets that may be explored such as Saturn and Jupiter. Venus is closer candidate however it has very harsh atmospheric conditions. High temperature, high density and corrosive nature of environment makes surface of the Venus challenging to survive.

Although Mars is the best candidate for human exploration, there are some challenges during two-way mission. Mars atmosphere has a density of 0.014 *kg=m*<sup>3</sup> and pressure of 610 *Pa* at the surface level. Low atmospheric density indicates that if Mars have had liquid water on its surface, it would have been evaporated immediately. In other words, the atmospheric pressure of Mars should be increased in

order to capture water molecules in atmosphere (at least 6.25 *kPa* means Armstrong limit). Armstrong limit is also critical for human body resistance. At this limit, breathable oxygen cannot be delivered to body more than a few minutes. Body fluids such as saliva, urine, tears, and alveoli in the lungs would boil away without a special pressurized body suit. Astronauts on Mars can live at this pressure level with a full-body pressure suit.

Human spaceflight to Red Planet takes at least 18 months with at least six months stay on the surface. Therefore, Martian air vehicles are needed both for surface operations and to return to the Earth. Thin atmosphere (due to low Reynolds number) indicates that only air vehicles such as micro helicopters or gliders can operate on Mars's atmosphere. However, these vehicles can only be used for observation purposes and are not feasible for transportation of large payloads. An advanced propulsion system is needed to fulfill mission requirements for long term two-way missions.

Current propulsion systems are quite expensive and technologically not feasible to fulfill two-way mission. In-situ Resources Utilization (ISRU) technologies are required for low cost and feasible propulsion systems. Both air breathing and rocket engines can be used as an ISRU based system. However, airbreathing engines need extremely large inlet areas due to the low atmospheric pressure of the Red Planet. Moreover, the condensed phase combustion products make the turbojet engine impractical. All these circumstances make rocket propulsion systems more practical for Martian operations [2].

Therefore, this chapter proposes a novel propulsion system for Mars Ascent Vehicles. Classical hybrid rocket motor configuration is tested as the propulsion system. The concept is supported by practical motor tests. Hybrid motor uses paraffin wax as the fuel binder and metallic powder as the additive. Aluminum and magnesium are mixed with paraffin as fuel additive. Experiments uses *CO*2*=N*2*O* mixture as the oxidizer agent. Combustion process as follows; Nitrous oxide reacts with paraffin and melts the metal oxide layer. Then carbon dioxide burn with metallic additive.

There are several factors to use *Paraffin=Metal=CO*2*=N*2*O* propellant combination. 96% of Mars atmosphere involves the *CO*<sup>2</sup> which is quite promising for in-situ missions. Carbon dioxide is known as a natural combustion product. It has also fire extinguisher feature. However, *CO*<sup>2</sup> can only burn with metallic powders. Therefore, *Metal=CO*<sup>2</sup> combustion releases significant amount of energy which is quite practical for sustainable Martian operations. *CO*<sup>2</sup> is self-pressurizing agent that removes the need for an additional pressurizing system in the rocket.

Experimental results show that magnesium has better ignition capability than alumiinum. *CO*<sup>2</sup> combustion is achieved up to 75% by mass in the oxidizer mixture. Adiabatic flame temperature of the motor is the key parameter for sustainable combustion of the carbon dioxide. Motor ignition quenches below the 1700 K. This is considered as the maximum flammability limit of the *CO*<sup>2</sup> with paraffin based hybrid motor.

Furthermore, this chapter provides brief knowledge related to hybrid propulsion systems. Hybrids offer safe, reliable, non-hazardous and cost-effective system compared to both liquid and solid systems [3, 4]. Fundamentals of hybrids briefly explained. Common propellant types are stated for hybrid motors. Details of *CO*<sup>2</sup> based experiments are presented accompanied with fuel grain manufacturing, oxidizer mixing process and combustion characteristics.

### **2. Hybrid rocket propulsion fundamentals**

This section provides fundamental information related to hybrid rocket propulsion. Hybrid rockets provide safety, reliability and environment friendly

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

manufacturing compared to other conventional rockets. It stores propellants in separate phases as in **Figure 1** [5]. Usually, the oxidizer is in liquid (or gaseous) phase and fuel is in solid phase.

The oxidizer is driven through a main valve into the solid fuel. An injector is used in order to control the oxidizer flow rate. Brass is the common material for the injector due to material properties. Brass is quite durable and resistant to high temperatures. The pressure of the oxidizer is regulated by pressurization system. Helium gas is mostly used as an pressurizing agent. There is an igniter which is placed near the injector manifold for initiating the ignition inside the motor. Igniter compounds moslty consist of potassium nitrate based solid fuels. However, larger rocket motors requires additional igniter motors as presented in [2]. Hybrid rocket motors mostly use circular port grain design. Circular port grains provide easy manufacturing and high regression rates [6]. Combustion occurs inside the grain port by the oxidizer flow and eventually exits through the nozzle. Nozzle throat diameter is designed in order to provide particular chamber pressure inside the motor. Graphite is mostly used for the nozzle manufacturing.

#### **2.1 Hybrid propulsion advantageous**

In the solid rocket motors, oxidizer and fuel are mixed as single solid phase. Combustion occurs by heating the solid fuel grain to reach the ignition temperature. Ignition of solid fuel cannot be stopped when it's started thus causes explosive danger. Thrust cannot be adjusted in solid motors that additional control systems are required for the rocket. Liquid motors keep oxidizer and fuel in separate tanks that combustion occurs by mixing propellant in a combustion chamber. Intimate mixture of propellant in a single chamber may cause explosion hazard. Propellant storage also requires exceptional cooling system in pumps, feed system and nozzle.

Hybrid rocket motors, however keeps oxidizer in liquid phase and fuel in solid phase [4]. Oxidizer delivery system (by using single oxidizer tank) reduce the complexity of plumbing compared to liquids. Thus, the ignition can be throttled by a main valve unlike solid motors. Hybrids have inert solid fuel grain thus; the grain manufacturing is safer than solid motors. Besides, it is easy to cast metallic additives in fuel grains to improve the combustion performance. Thus, operation feasibility at low temperatures, long oxidizer storage capability with non-hazardous manufacturing make the hybrids more practical for Mars missions.

#### **2.2 Propellant evaluation in hybrid rockets**

Classical hybrid rockets commonly use polymeric fuels such as HTPB (Hydroxyl-Terminated Polybutadiene), HTPE (Hydroxyl-terminated polyether) and PE (Polyethylene). In hybrid motors, a turbulent boundary layer is formed by

**Figure 1.** *Paraffin wax liquid entertainment mechanism.*

oxidizer injection over the polymeric fuel surface [6]. Thus, the diffusion flame occurs during the ignition at the boundary layer. Diffusion flame is transported on the surface by oxidizer flow. Radiation and convection heat transfer play an important role during this process. Thereby, vaporized fuel on the grain surface reacts with atomized liquid oxidizer causes "blocking effect" at the wall. This blocking effect limits the burn rate of the motor [7].

Paraffin wax is another fuel commonly used in hybrid motors [4]. Paraffin enables 3 to 5 times higher regression rate at the same oxidizer mass flux compared to classical polymeric fuels [6]. Burning paraffin fuel produces a liquid layer over the fuel grain with low viscosity and low surface tension. The liquid melt layer consists of liquified paraffin fuel droplets. This layer merges with the oxidizer flow becomes hydrodynamically unstable in the fuel port. Unstable ignition creates an instability; it lift-offs paraffin fuel droplets from the grain surface that foster the mass transfer rate of fuels into the oxidizer gas flow. This is called as "liquid entrainment mass transfer mechanism" can be seen in **Figure 2** [7]. This additional mass transfer mechanism increases the regression rate of hybrid motor combustion.

Hybrid rockets commonly use *N*2*O*<sup>4</sup> (dinitrogen tetraoxide), *H*2*O*<sup>2</sup> (hydrogen peroxide), gaseous ð Þ *GOX* or liquid oxygen ð Þ *LOX* and *N*2*O* (nitrous oxide) as the oxidizer [8]. Nitrogen tetraoxide is storable high-density oxidizer that is used in early launch vehicles. *N*2*O*<sup>4</sup> provides moderate *Isp* with performance additives. However, it's a high toxic chemical. Hydrogen peroxide is also storable and highdensity agent. *H*2*O*<sup>2</sup> is an aggressive chemical at high concentrations. It has leaning to self-decompose thus causes detonation hazard. Typical rocket applications use over 80% concentration levels of *H*2*O*<sup>2</sup> that makes the distillation and handling quite critical for human skin. Liquid oxygen is the most common high-performance oxidizer in the rocket industry. *LOX* is highly stable due to diatomic oxygen bond and provides high specific impulse. It provides lower oxidizer to fuel ratio that reduces the fraction of oxidizer used in the propulsion system. It is also costeffective compound. However, *LOX* is cryogenic material with boiling temperature of 90 K. Cryogenic nature of makes it challenging during surface operations on the Mars. Moreover, liquid oxygen is not self-pressurizing agent due to its low density needs an additional pressurization system via Helium or Nitrogen. This increases complexity and cost of a possible Martian rocket.

Nitrous oxide is another agent has been used mostly in small rocket systems. Nitrous has self-pressurizing capability at saturated liquid state [2]. Selfpressurization eliminates the need of pressurizing system to feed the oxidizer. Thus, it reduces the complexity, weight and cost of the propulsion system.

**Figure 2.** *Paraffin wax liquid entertainment mechanism.*

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

Self-pressurization feature also makes nitrous an efficient candidate for Mars environment. *N*2*O* is non-toxic and easy to handle compared to both *N*2*O*4, *H*2*O*<sup>2</sup> and *LOX*. Nitrous oxide also creates a highly exothermic decomposition reaction during the combustion. Therefore, it provides stable and efficient ignition in the rocket systems. Nitrous oxide in liquid phase is quite safe and easy to store at room temperature. Highly storable feature makes launch operations quite easy compared to other oxidizer options. Also, its readily available in chemical industry. On the other hand, reduced specific impulse, low density at higher temperatures and strong dependence of the temperature are among several disadvantages of *N*2*O*. It should be noted that nitrous has positive heat of formation. Thus, selfdecomposition of "vapor phase" nitrous in feed lines, oxidizer tank and combustion chamber result disruptive damage [9].
