**3. Propulsion options**

High-thrust chemical propulsion, using oxygen/hydrogen (O2/H2) rocket engines is a natural choice [12]. If indeed water were available on the Martian moons, it would make sense to capitalize on that water resource.

Electric propulsion systems with either ion or Hall thrusters are potential options. Xenon or other inert gases are the typical choice for such thrusters. Using hydrogen as an electric propulsion propellant has also been proposed [12]. However, the hydrogen propellant option is a far term prospect [12].

Mass scaling equations were developed for the O2/H2 and the nuclear electric propulsion (NEP) systems [12].

#### **Figure 2.**

*Mission options and delta-V—Deimos, using low thrust (blue) and high thrust (orange).*

**57**

*Assessing Propulsion and Transportation Issues with Mars' Moons*

*DOI: http://dx.doi.org/10.5772/intechopen.93148*

**3.1 Advanced propulsion options**

*Orbital transfer—high-altitude plane change, 100,000 km.*

*3.1.1 Chemical propulsion OTV sizing*

m, p = usable propellant mass (kg); and a, fixed = chemical OTV fixed mass (kg).

are presented.

**Figure 4.**

where

propellant mass);

*3.1.2 NEP OTV sizing*

Several advanced propulsion options for lunar base construction and industrialization were investigated. They include nuclear electric propulsion options, lunar base design options, propellant industrialization, and outer planet mining with associated outer planet moon bases. Chemical propulsion and nuclear electric propulsion (NEP) for Earth-Moon orbital transfer vehicles (OTVs) were assessed. Design parameters, vehicle mass scaling equations, and summaries of these analyses

In sizing the chemical propulsion OTVs, a vehicle mass scaling equation is used [12]:

m, dry, stage = the stage dry mass, including residual propellant (kg); m, dry, coefficient = the B mass coefficient (kg of tank mass/kg of usable

system masses. The Martian moon OTVs were single-stage vehicles.

The chemical propulsion OTVs had a B coefficient of 0.4. The fixed mass was 500 kg. The fixed mass includes guidance systems, adapters, and reaction control

The NEP OTV mass and trip time were estimated based on the power system and the propulsion system design [11]. The following dry mass scaling equation was used [11]:

*m dry stage m dry coefficient m p a fixed* , , , , \*, , = + (1)

**Figure 3.** *Orbital transfer at low altitude—Phobos and Deimos altitude.*

*Assessing Propulsion and Transportation Issues with Mars' Moons DOI: http://dx.doi.org/10.5772/intechopen.93148*

*Mars Exploration - A Step Forward*

propulsion (NEP) systems [12].

High-thrust chemical propulsion, using oxygen/hydrogen (O2/H2) rocket engines is a natural choice [12]. If indeed water were available on the Martian

Electric propulsion systems with either ion or Hall thrusters are potential options. Xenon or other inert gases are the typical choice for such thrusters. Using hydrogen as an electric propulsion propellant has also been proposed [12]. However,

Mass scaling equations were developed for the O2/H2 and the nuclear electric

moons, it would make sense to capitalize on that water resource.

*Mission options and delta-V—Deimos, using low thrust (blue) and high thrust (orange).*

the hydrogen propellant option is a far term prospect [12].

**3. Propulsion options**

**56**

**Figure 3.**

**Figure 2.**

*Orbital transfer at low altitude—Phobos and Deimos altitude.*

**Figure 4.** *Orbital transfer—high-altitude plane change, 100,000 km.*

#### **3.1 Advanced propulsion options**

Several advanced propulsion options for lunar base construction and industrialization were investigated. They include nuclear electric propulsion options, lunar base design options, propellant industrialization, and outer planet mining with associated outer planet moon bases. Chemical propulsion and nuclear electric propulsion (NEP) for Earth-Moon orbital transfer vehicles (OTVs) were assessed. Design parameters, vehicle mass scaling equations, and summaries of these analyses are presented.

#### *3.1.1 Chemical propulsion OTV sizing*

In sizing the chemical propulsion OTVs, a vehicle mass scaling equation is used [12]:

$$m, dr\gamma, \text{stage} = m, dr\gamma, \text{coefficient } "\,\, m, p+a, \text{fixed} \tag{1}$$

where

m, dry, stage = the stage dry mass, including residual propellant (kg); m, dry, coefficient = the B mass coefficient (kg of tank mass/kg of usable propellant mass);

m, p = usable propellant mass (kg); and

a, fixed = chemical OTV fixed mass (kg).

The chemical propulsion OTVs had a B coefficient of 0.4. The fixed mass was 500 kg. The fixed mass includes guidance systems, adapters, and reaction control system masses. The Martian moon OTVs were single-stage vehicles.

#### *3.1.2 NEP OTV sizing*

The NEP OTV mass and trip time were estimated based on the power system and the propulsion system design [11]. The following dry mass scaling equation was used [11]:

where

m, dry, stage, NEP = NEP dry mass (kg); alpha = NEP reactor specific mass (kg/kWe); P = NEP power level (kWe); 0.05 = tankage mass coefficient (kg/kg m, p); m, p = NEP usable propellant mass (kg); and m, fixed = NEP fixed mass (kg).

The OTV sizing was conducted for a wide range of power levels: 0.5–30 MWe. Three nuclear reactor-specific masses were used: 10, 20, and 40 kg/kWe (kilograms per kilowatt, electric). The OTV propulsion fixed mass, apart from and in addition to the reactor mass, was 20 MT, and the propellant tankage mass was 5% of the mass of the required propellant.

The Isp and efficiency of the electric propulsion systems were 5000 seconds with overall thruster propulsion efficiencies of 50% for each design. These design points are typical of advanced designs of either magnetoplasmadynamic (MPD) or pulse inductive thrusters (PIT). While hydrogen is suggested for both propulsion system thrusters, the possibilities of the higher Isp option using inert gases (xenon, krypton, etc.) are also viable. The low-thrust OTV delta-V value varied based on the destination of the Martian moon missions.

## **4. Mission effectiveness**

#### **4.1 Phobos and Deimos payload missions**

**Figures 5** and **6** depict the Phobos and Deimos O2/H2 propulsion system initial masses. For the 50 MT payload case for Phobos, the OTV initial mass is 141 MT. Nearly the same OTV mass is needed to perform the Phobos to LMO and Phobos to 100,000 km. For Deimos, the highest OTV mass is 256 MT.

The payload mass cases presented range from 1, 10, 20 to 50 MT. If the payload mass is less than 10 MT, the O2/H2 OTV mass is very small. If small 1 MT payloads must be sent quickly from one orbit to another, the O2/H2 OTV is an excellent choice; the propellant mass of the chemical propulsion system is very low compared to the NEP propellant mass. Alternatively, if five 10 MT payloads can be manifested together, the NEP OTV has a significant propellant mass advantage over the O2/H2 OTV.

**Figures 7**–**9** present the round-trip mission trip times for the NEP vehicles for 1, 10, and 50 MT, respectively. The NEP trip times are many days long: for a 5 MWe NEP OTV with a 50 MT payload, the trip time for Phobos to LMO is 55 days, whereas the chemical propulsion trip times are less than 1 day. However, the benefits of reduced NEP propellant resupply mass are quite significant.

With NEP OTVs, the 10 MWe power levels provide the shortest trip time; however, if the payload can be delivered more slowly, the 1 MWe power level allows a very large propellant mass savings over the higher 10 MWe power level. The NEP propellant mass savings for large payloads are a critical part of any sustainable architecture. The propellant mass savings are noted in the succeeding sections.

Fast transfers of critical items under 1 MT are best accomplished with O2/H2 OTV propulsion. There may be a critical need for the delivery of medical supplies;

**59**

**Figure 6.**

**Figure 5.**

*Assessing Propulsion and Transportation Issues with Mars' Moons*

also the delivery of space parts or a repair crew may be needed. The O2/H2 OTV

**Figures 10**–**13** compared the propellant masses for the O2/H2 system with several NEP systems. For the NEP cases, power levels of 0.5–10 MWe are shown. **Figures 10** and **11** present the Phobos and Deimos cases for 10 MT payloads, and the 50 MT payload cases are shown in **Figures 12** and **13**. In **Figure 10**, for a 10 MT payload, the Phobos to LMO NEP cases will allow large propellant savings over the O2/H2 OTV for NEP OTV power levels of less than 5 MWe. In the Deimos case, shown in **Figure 11**, the NEP OTV provides significant propellant mass savings over O2/H2 with power levels up to 10 MWe. The **Figure 12** data for Phobos to LMO with

would be best suited for these small 1 MT payloads.

*Deimos OTV initial masses—O2/H2 propulsion.*

*DOI: http://dx.doi.org/10.5772/intechopen.93148*

*Phobos OTV initial masses—O2/H2 propulsion.*

#### *Assessing Propulsion and Transportation Issues with Mars' Moons DOI: http://dx.doi.org/10.5772/intechopen.93148*

*Mars Exploration - A Step Forward*

P = NEP power level (kWe);

mass of the required propellant.

**4. Mission effectiveness**

m, fixed = NEP fixed mass (kg).

destination of the Martian moon missions.

**4.1 Phobos and Deimos payload missions**

100,000 km. For Deimos, the highest OTV mass is 256 MT.

m, dry, stage, NEP = NEP dry mass (kg); alpha = NEP reactor specific mass (kg/kWe);

0.05 = tankage mass coefficient (kg/kg m, p); m, p = NEP usable propellant mass (kg); and

where

*m dry stage NEP alpha P m p m fixed* ,, , = \*+ + 0.05 , , (2)

The OTV sizing was conducted for a wide range of power levels: 0.5–30 MWe. Three nuclear reactor-specific masses were used: 10, 20, and 40 kg/kWe (kilograms per kilowatt, electric). The OTV propulsion fixed mass, apart from and in addition to the reactor mass, was 20 MT, and the propellant tankage mass was 5% of the

The Isp and efficiency of the electric propulsion systems were 5000 seconds with overall thruster propulsion efficiencies of 50% for each design. These design points are typical of advanced designs of either magnetoplasmadynamic (MPD) or pulse inductive thrusters (PIT). While hydrogen is suggested for both propulsion system thrusters, the possibilities of the higher Isp option using inert gases (xenon, krypton, etc.) are also viable. The low-thrust OTV delta-V value varied based on the

**Figures 5** and **6** depict the Phobos and Deimos O2/H2 propulsion system initial masses. For the 50 MT payload case for Phobos, the OTV initial mass is 141 MT. Nearly the same OTV mass is needed to perform the Phobos to LMO and Phobos to

The payload mass cases presented range from 1, 10, 20 to 50 MT. If the payload mass is less than 10 MT, the O2/H2 OTV mass is very small. If small 1 MT payloads must be sent quickly from one orbit to another, the O2/H2 OTV is an excellent choice; the propellant mass of the chemical propulsion system is very low compared to the NEP propellant mass. Alternatively, if five 10 MT payloads can be manifested together, the NEP OTV has a significant propellant mass advantage over the

**Figures 7**–**9** present the round-trip mission trip times for the NEP vehicles for 1, 10, and 50 MT, respectively. The NEP trip times are many days long: for a 5 MWe NEP OTV with a 50 MT payload, the trip time for Phobos to LMO is 55 days, whereas the chemical propulsion trip times are less than 1 day. However, the benefits of reduced NEP propellant resupply mass are quite

With NEP OTVs, the 10 MWe power levels provide the shortest trip time; however, if the payload can be delivered more slowly, the 1 MWe power level allows a very large propellant mass savings over the higher 10 MWe power level. The NEP propellant mass savings for large payloads are a critical part of any sustainable architecture. The propellant mass savings are noted in the succeeding

Fast transfers of critical items under 1 MT are best accomplished with O2/H2 OTV propulsion. There may be a critical need for the delivery of medical supplies;

**58**

O2/H2 OTV.

significant.

sections.

#### **Figure 5.** *Phobos OTV initial masses—O2/H2 propulsion.*

**Figure 6.** *Deimos OTV initial masses—O2/H2 propulsion.*

also the delivery of space parts or a repair crew may be needed. The O2/H2 OTV would be best suited for these small 1 MT payloads.

**Figures 10**–**13** compared the propellant masses for the O2/H2 system with several NEP systems. For the NEP cases, power levels of 0.5–10 MWe are shown. **Figures 10** and **11** present the Phobos and Deimos cases for 10 MT payloads, and the 50 MT payload cases are shown in **Figures 12** and **13**. In **Figure 10**, for a 10 MT payload, the Phobos to LMO NEP cases will allow large propellant savings over the O2/H2 OTV for NEP OTV power levels of less than 5 MWe. In the Deimos case, shown in **Figure 11**, the NEP OTV provides significant propellant mass savings over O2/H2 with power levels up to 10 MWe. The **Figure 12** data for Phobos to LMO with

**Figure 7.** *NEP OTV trip time, 1 MT payload: Phobos to LMO.*

**Figure 8.** *NEP OTV trip time, 10 MT payload: Phobos to LMO.*

a 50 MT payload shows very large NEP propellant mass reductions over O2/H2 for the 10 MWe power level; for a 5 MWe power level, the propellant mass reduction was from 57 to 10 MT. Similarly, in **Figure 13**, the Deimos to LMO cases show very significant propellant mass benefits, reducing the propellant needed by a factor of 6–10 or more over O2/H2. In nearly all cases, the NEP systems allow large propellant mass reductions. For large mission architectures over a long-term Mars project, the mass reductions can be as high as a factor of 5–10 over O2/H2 systems.

### **4.2 Mars lander options**

Past studies of Mars landers have included an innovative single stage to orbit (SSTO) design [15]. The Mars Base Camp mission suggested an aerospacecraft that

**61**

**Figure 10.**

*payload mass.*

**Figure 9.**

wrested from the moon's regolith.

*Assessing Propulsion and Transportation Issues with Mars' Moons*

would carry an astronaut crew to the surface of Mars and return to orbit, all with a single stage. The Mars sortie vehicle would be refueled with oxygen and hydrogen propellants created from in situ water resources from the Martian moons. A water electrolysis factory would be delivered to one of the moons and the water would be

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Phobos to LMO, 10 MT* 

In Reference [15], the Mars sortie vehicle was designed to use 80 MT of O2/H2

**Figure 14** presents the initial mass and propellant mass for a Mars sortie vehicle. The dry mass fraction, B, is varied from 0.1 to 0.25. In the Reference [15] analysis, the 80 MT propellant load would require a somewhat optimistic B fraction of less than 0.10. Using a B fraction of 0.2, the required propellant mass

propellant. The initial mass would be approximately 108 MT.

*DOI: http://dx.doi.org/10.5772/intechopen.93148*

*NEP OTV trip time, 50 MT payload: Phobos to LMO.*

#### *Assessing Propulsion and Transportation Issues with Mars' Moons DOI: http://dx.doi.org/10.5772/intechopen.93148*

#### **Figure 10.**

*Mars Exploration - A Step Forward*

**60**

**Figure 8.**

**Figure 7.**

*NEP OTV trip time, 1 MT payload: Phobos to LMO.*

*NEP OTV trip time, 10 MT payload: Phobos to LMO.*

**4.2 Mars lander options**

a 50 MT payload shows very large NEP propellant mass reductions over O2/H2 for the 10 MWe power level; for a 5 MWe power level, the propellant mass reduction was from 57 to 10 MT. Similarly, in **Figure 13**, the Deimos to LMO cases show very significant propellant mass benefits, reducing the propellant needed by a factor of 6–10 or more over O2/H2. In nearly all cases, the NEP systems allow large propellant mass reductions. For large mission architectures over a long-term Mars project, the

Past studies of Mars landers have included an innovative single stage to orbit (SSTO) design [15]. The Mars Base Camp mission suggested an aerospacecraft that

mass reductions can be as high as a factor of 5–10 over O2/H2 systems.

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Phobos to LMO, 10 MT payload mass.*

would carry an astronaut crew to the surface of Mars and return to orbit, all with a single stage. The Mars sortie vehicle would be refueled with oxygen and hydrogen propellants created from in situ water resources from the Martian moons. A water electrolysis factory would be delivered to one of the moons and the water would be wrested from the moon's regolith.

In Reference [15], the Mars sortie vehicle was designed to use 80 MT of O2/H2 propellant. The initial mass would be approximately 108 MT.

**Figure 14** presents the initial mass and propellant mass for a Mars sortie vehicle. The dry mass fraction, B, is varied from 0.1 to 0.25. In the Reference [15] analysis, the 80 MT propellant load would require a somewhat optimistic B fraction of less than 0.10. Using a B fraction of 0.2, the required propellant mass

#### **Figure 11.**

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Deimos to LMO, 10 MT payload mass.*

#### **Figure 12.**

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Phobos to LMO, 50 MT payload mass.*

is nearly 200 MT. Therefore five, 40 MT ISRU water resupply flights would be required to support any Mars sortie vehicles.

Using electric propulsion for the resupply flights would enhance the overall architecture, by significantly reducing the total propellant mass needed for the sortie vehicle refueling. Five NEP resupply flights from Phobos would require 50 MT, whereas nearly 300 MT (approximately 6 times the mass) of O2/H2 propellant are needed to transport that propellant to LMO. Many propellant deliver benefits are also gained at lower NEP power levels.

**63**

**5. Concluding remarks**

**Figure 14.**

**Figure 13.**

*payload mass.*

radars and other high-energy science instruments.

*Mars sortie vehicle initial mass and propellant masses: SSTO capability.*

(compared to the propellant for O2/H2 OTVs).

Electric propulsion offers the ability to transfer large payloads between the Martian moons and in Mars orbit space over O2/H2 propulsion. The benefits of electric propulsion are not only in the reduction of propellant masses, but the capability of the high-power reactor system to perform unique science investigations, using

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Deimos to LMO, 50 MT* 

Mining the moons will require specialized factories and processes. The extremely low gravity on the Martian moons will be a challenge in controlling dust, anchoring the spacecraft factory and controlling processes. An artificial

The NEP systems have a flexible design and can allow many payloads to be manifested together, reducing the overall propulsion architecture. High inclination Mars orbits can be more easily accessed with NEP OTVs with small amount of propellant

*Assessing Propulsion and Transportation Issues with Mars' Moons*

*DOI: http://dx.doi.org/10.5772/intechopen.93148*

#### *Assessing Propulsion and Transportation Issues with Mars' Moons DOI: http://dx.doi.org/10.5772/intechopen.93148*

#### **Figure 13.**

*Mars Exploration - A Step Forward*

**62**

**Figure 12.**

**Figure 11.**

*payload mass.*

*payload mass.*

is nearly 200 MT. Therefore five, 40 MT ISRU water resupply flights would be

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Phobos to LMO, 50 MT* 

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Deimos to LMO, 10 MT* 

Using electric propulsion for the resupply flights would enhance the overall architecture, by significantly reducing the total propellant mass needed for the sortie vehicle refueling. Five NEP resupply flights from Phobos would require 50 MT, whereas nearly 300 MT (approximately 6 times the mass) of O2/H2 propellant are needed to transport that propellant to LMO. Many propellant deliver benefits

required to support any Mars sortie vehicles.

are also gained at lower NEP power levels.

*Resupply propellant mass and round-trip time for O2/H2 and Xe Ion NEP OTVs—Deimos to LMO, 50 MT payload mass.*

#### **Figure 14.**

*Mars sortie vehicle initial mass and propellant masses: SSTO capability.*

## **5. Concluding remarks**

Electric propulsion offers the ability to transfer large payloads between the Martian moons and in Mars orbit space over O2/H2 propulsion. The benefits of electric propulsion are not only in the reduction of propellant masses, but the capability of the high-power reactor system to perform unique science investigations, using radars and other high-energy science instruments.

The NEP systems have a flexible design and can allow many payloads to be manifested together, reducing the overall propulsion architecture. High inclination Mars orbits can be more easily accessed with NEP OTVs with small amount of propellant (compared to the propellant for O2/H2 OTVs).

Mining the moons will require specialized factories and processes. The extremely low gravity on the Martian moons will be a challenge in controlling dust, anchoring the spacecraft factory and controlling processes. An artificial gravity factory will likely be needed to maintain the water and propellant processing quality.
