4 | Moon Missions

I. Introduction

Humanity has long stared at the pale white Moon and dreamed of starting a civilization there.

Starting from the 50th anniversary of the first Apollo landing, we will assume there will be an

international collaborative effort to establish a permanent lunar foothold. Throughout this chapter,

we will discuss our missions to the Moon, how we plan to stay, and how we will develop industries

necessary to fuel humanity’s plans to Mars.

II. LaGrange point and Halo orbits

As per the Purdue-Aldrin specifications, missions to both L1 and L2 are to be examined and

communications relay network from Earth to L2 established.

A. L1

Before the trajectory to L1 is calculated, we first solve for the location of L1 with respect to the

lunar orbit. The calculations for this are shown in Appendix (XX). We then calculate for the Delta

V and time of flight required to place an XM1 (1st generation exploration module) on a trajectory

to L1 from Earth. A low energy transfer along a precalculated stable manifold is used to execute

the transfer. The disadvantage of using this low energy transfer, as is the disadvantage of most low

energy transfers, is that the length of the TOF is significantly longer. The orbital characteristic of

the transfer along the manifold are shown in Table XX below:

Table 4.1: Transfer characteristics of the stable manifold from Earth to L1 .

Characteristic Value [unit]

Delta V 3.2249 [km/s]

TOF 76.4852 [days]

B. L2 – Halo Orbit

Second, we look at the LaGrange point L2. We decide to use the XM1 initially to be at the L2

point also as a means of the communication relay network from Earth to L2. To do this, we place

the XM1 in a halo orbit about L2. In order to get to the halo orbit, a second precalculated stable

manifold is used to get to L2. Once at L2, a burn is executed to get the module into the halo orbit.

The orbital characteristic of the transfer along the manifold and the Delta V requirements of the

halo orbit are shown in Table 4.2 below:

Table 4.2: Transfer characteristics of the stable manifold from Earth to L1 .

Characteristic Value [unit]

Delta V to L2 3.0957 [km/s]

Delta V to L2 Halo 0.1010 [km/s]

TOF 90.9461 [days]

III. Refueling Station

As per the Purdue – Aldrin specifications, we establish a refueling depot in a geocentric orbit

in cis-lunar space in order to provide a means of supplying fuel to modules should the need arise.

To pick an orbit for the refueling station, we consider the criteria for the vehicle to be in orbit. A

high velocity, which keeps the period small, will eliminate the incovenience of a long wait for an

orbital return. However, the altitude should also be high enough to eliminate any inconveniences

in the form of other orbital vehicles. The orbital characteristics of the refueling station orbit are

shown in Table 4.3 below:

Table 4.3: Transfer characteristics of the stable manifold from Earth to L1 .

Characteristic Value [unit]

Altitude 17000 [km]

Velocity 4.082 [km/s]

Period 10 [hours]

IV. Cargo Modules

To place the cargo module in low lunar orbit (LLO), we consider a low thrust transfer from

Earth to the lunar vicinity. The cargo vehicle leaves low earth orbit (LEO) at a 200km altitude.

Due to the degree of similarity of the missions, the low thrust solution from Project Artemis -

2014, is modified in order to obtain mission specific values for the Purdue-Aldrin mission.

Detailed explanation of the method and solution can be found in Apeendix 4.3. The mission

characteristics are shown in Table 4.3 below:

Table 4.4: Transfer characteristics of cargo mission to LLO .

Characteristic Value [unit]

Mass 225.86 Mg

Total approx. propellant cost 118.1 Mg

Total TOF 402 days

V. Moon Base Human Factors Requirements

Establishing a permanent human presence on the Moon will be an important step towards

the eventual goal of colonizing Mars. The lessons learned by establishing this settlement will be

invaluable in the process of further refining our design for a Mars colony. Thus, the moon base

will employ early stages of the same systems that we designed for Mars. Crews on the Moon will

live in the XM2’s, (Exploration Module - 2) which are early variants of the XM3 habitation

modules that will be used on Phobos and Mars.

There will be three bases established on the Moon. Each will be made up of three

habitation modules (XM-2’s), totaling eighteen people per base and 54 people living on the lunar

surface. In accordance with the mission requirements, one base will be located on the near side

of the moon, one on the far side, and a third in the Shackelton Crater.

Human Factors needs for the settlement on the Moon are the same as for the colony on Mars.

The main requirments will be food and water. The food requirments will be the same 2600

calories allotted for the colony on Mars. The water will be used for drinking and hygiene, as well

as for oxygen production via electrolysis. The Nitrogen supply is used to dilute the oxygen in the

habitation module’s atmosphere for health and fire safety reasons. There is also a supply of

backup oxygen stored in liquid form that can provide a breathable atmosphere for the crew for 60

days in the event that the oxygen production systems fail. The following table shows the human

factors needs for each base for one year.

Table 4.5: Moon Base Human Factors Requirments (Crew of 18)

Total Mass, Mg Total Volume, m3

Food 14.34 18.03

Water 4.2 4.2

Nitrogen 0.014 0.017

Backup Oxygen 0.885 0.7762

These numbers are based on the assumption that all needed supplies are sent from Earth.

There is hope that the lunar bases will be able to harvest water from Shackleton Crater, and the

aeroponic farming systems being developed for Mars could be used at the moon as well.

However, we are going to use the Moon as a testing ground for these tehcnologies, so it is better

to prepare in such a way that the lunar bases do not rely on them.

The life support and water recovery systems will be same ones planned for use on the Mars

colony. Water recovery rate for this system is projected to be 91%. More information on life

support and water systems can be found in Appendices X and Y. These systems, as well as other

human factors needs such as cooking and cleaning will require power. The following table shows

the maximum power requirments for each XM2, as well as a total for an entire lunar base (three

modules).

Table 4.6: Max Power Requirements for a Lunar Base

Max Power, kW

Crew Quarters (x3) 19.77

Water Systems/Life Support (x3) 29.4

Base Total 147.51

Additional human factors considerations are radiation shielding and mitigating the effects of a

low-gravity environment. Radiation shielding will be provided by piling lunar regolith in and

around the habitation modules. To mitigate the effects of spending long periods of time in low-

gravity, the crew will spend time every day exercising in a small centrifuge that provides Earth

levels of gravity. Neither of these topics has been addressed thoroughly as they are beyond the

scope of the current project, but a more in-depth discussion of radiation effects can be found in

Appendix X and the effects of low-gravity are discussed in Appendix Y.

VI. Lunar In-Situ Propellant Production

Interplanetary missions to Mars require immense sums of propellant in order to generate the

velocities necessary to leave Earth and go to Mars. If humanity wants to develop a colony on the

red planet, we will need to have the ability to not only send large amount of payload to Mars, but

frequently too. Therefore the motivation of lunar in-situ propellant production is to provide access

to a long term sustainable supply of propellant to power the spaceships of tomorrow. This section

will assume the lunar colonists have chosen Shackleton Crater as a colony site in order to explore

possibility of in-situ propellant production.

A. Shackleton Crater Regolith Properties

In this section, we will examine the property of regolith at Shackleton Crater in order to provide

the core assumptions used for our ISRU analysis. The following tables details the values used

throughout the analysis:

Table 4.7: Regolith Properties at Shackleton Crater

B.

Propellant Considerations for In-Situ Propellant Production

We need to first evaluate what are viable propellants we can produce using lunar resources.

Three main propellants have been selected for study: liquid hydrogen & liquid oxygen, methane

& liquid oxygen, and silane & liquid oxygen.

Hydrogen (H2) is one of the most efficient chemical fuels. Liquid hydrogen rocket engines have

been flowed very successfully and reliably in the past few decades, so the technology is

Variable Value

Ice by Weight 6.5%

Density of Regolith 1.7 Mg/m3

Area of Minable Regolith 346 km2

Depth of Minable Regolith 1 m

proven. It can also be produced using water extracted from the Moon. While liquid hydrogen does

have problems with boiling off currently, we will assume zero boil off technology will be available

by the time of this mission.

Methane (CH4) serves as an alternative fuel to liquid hydrogen. Methane rocket engines have

recently seen a lot of research and development in the past few years. It methane rocket engines

will be flying missions by the end of this decade. While methane rockets provide less performance

than liquid hydrogen, it is much easier to storage due to its higher density and boiling temperature.

Despite these advantages, the Moon is very carbon poor and carbon makes up 89% of methane by

weight. Therefore we will need to obtain the carbon necessary for methane production elsewhere.

Silane (SiH4) serves as an alternative fuel to methane. Silane provides many of the same

advantages as methane. While silane provides less performance than methane, silane can be

produced on the Moon. Silicon can be harvested from the silicon rich regolith of the moon and

hydrogen can be produced from the ice rich regolith of Shackleton. However silane combustion is

not well understood and no silane rocket engine has ever been built and tested.

Oxygen (O2) can be found in abundant on the Moon in the form of various metal oxidizes.

Oxygen is also a major byproduct when producing hydrogen from water. Therefore the availability

of oxygen is of least concern.

Due to the limitation on producing methane on the Moon and the limitation of silane rocket

technology, liquid hydrogen remains the only viable option for ISRU propellant production. We

will proceed forward in the following sections with liquid hydrogen & liquid oxygen in mind as

our primary propellant.

C. Lunar ISRU Production Requirements and Analysis

In this section, we will examine the propellant production requirements we will need to meet so

we can provide all of the propellant necessary for the missions to Mars from the Moon. Throughout

the analysis, we are focused on the liquid hydrogen (LH2) mass requirement instead of the liquid

oxygen (LOX) mass requirement. Since we plan on producing LH2 and LOX through electrolysis,

LH2 is our limiting factor and will be our driving factor to determine how much power and land

we need to process to meet production quotas. To better understand our production goals, we need

to examine how much LH2 mass we need to provide every year for the various missions to Mars.

The following table outlines the total amount of LH2 mass required every launch window and the

amount of time between launches:

Table 4.7: LH2 Mass Requirements per Launch Window

The values above can be found in Appendix AK, AZ, and BX. Due to varying numbers of

vehicles launching each window, we will have fluxuation in LH2 mass requirements. These

fluxuation can causes peaks and lows in power demands due to varying production quota every

launch window. However we can eliminate these fluxuation by setting a constant yearly LH2

production rate. By ensuring our total cumulative output of LH2 exceeds our total cumulative

consumption of LH2 at all times, we can optimize the sizing of our propellant production plant.

Examining the figure below, we can visual represent our total cumulative production and

consumption of LH2 to aid our decision making:

Year XM3 Cycler CarLa HuLa LH2 (Mg) Period (Years)

2028 3 1 264.61 2

2030 2 111.18 2

2031 3 1 264.61 1

2032 1 55.59 1

2033 2 3 3 333.09 1

2035 3 1 264.61 2

2037 3 1 3 3 486.52 2

Figure 4.1: LH2 Production Rate vs LH2 Consumption Rate

From Figure 4.1, we can visually see that with an annual production rate of LH2 at 215 Mg, we

can provide enough propellant to all of our launch vehicles to Mars. Therefore, we are going to

size our propellant plant to produce 215 Mg of LH2 every year.

However before examining the sizing of our propellant plant, we need to make sure an annual

production rate of 215 Mg of LH2 is feasible and sustainable, given our ISRU background is at

Shackleton Crater. The table below details some relevance statistics with mining water at

Shackleton Crater with our desired consumption rate.

Table 4.8: Annual Statistics on Mining Water at Shackleton

Shackleton Value Units

Annual H2 Mass Mined 215 Mg

Annual Water Required 1935 Mg

Annual Regolith Required 29770 Mg

Total Amount of Water Available 38,270,000 Mg

Annual H2O Depletion Rate 0.00506 %

With an annual H2O depletion rate of less than 0.01%, we can safely say mining water at

Shackleton in order to produce propellant is a sustainable option.

D. Lunar ISRU Production Plant Overview

We will satisfy our propellant demands through the use of two processes: water extraction and

propellant production. In the water extraction process, we first extract the water from the regolith

by heating the regolith up the boiling point of water in a low pressure furnace. Next, we will

compress the water vapor to make it easier to condenser. Finally, we will cool the water vapor

through a condenser until it condenses into a liquid. A thermal regeneration cycle is added to

reduce overall system power requirement by using the energy tapped off from the cooling water

vapor. The following figure on the next page is a diagram of the water extraction process.

The next, the propellant production system uses the water produced from the water extraction

process to create our required propellant. We accomplish this by first running the liquid water

through a PEM electrolysis system that splits the water into hydrogen and oxygen gas. The gases

will then be filtered and separated. The separated oxygen gas will then be cooled through an

oxygen liquefier system which will liquefy the oxygen gas for storage. The hydrogen gas however

will have its pressure raised through a compressor first.

Figure 4.2: The complete Water Processing and Regeneration cycle

Due to the low boiling point temperature of hydrogen gas, it is advantageous to increase its

pressure in order to raise the boiling point temperature. A higher boiling point temperature will

mean less energy required to liquefy the gas. Once the pressure of the hydrogen gas is raised to

our desired conditions, we will finally cool it through a hydrogen liquefier to turn it into liquid

form for storage.

The helium loop in this system acts mainly as the coolant for the oxygen and hydrogen liquefier.

It will also have a radiator to remove the excess heat it absorbs through the liquefier system. We

can assume a radiator system can work to cool the helium to extremely low temperatures because

we are in a permanently shadowed region at the South Pole of the Moon. The following figure is

a diagram of the propellant production cycle.

Figure 4.3: The complete Propellant Production and Cooling cycle

Further in-depth analysis of the lunar ISRU production plant can be found on Appendix Z and

Appendix AA.

E. Lunar ISRU Production Plant Sizing

Since we need to produce at 215 Mg of H2 annually, we can size our lunar ISRU production

plant to a maximum yield of 215 Mg of H2 per year. The following table describes the

specifications of the propellant production plant:

Table 4.9: Lunar ISRU Production Plant Specifications

System Power

(MW)

Mass

(Mg)

Volume

(m3)

Water Production

Plant 2.04 3.72 5.35

Propellant

Production Plant 1.56 3.56 6.47∙103

Total ISRU Plant 3.60 7.28 6.47∙103

Based on these values, the power and mass requirements are manageable at 3.6 MW and 7.28

Mg respectively. The system volume is massive because it includes the volume necessary for the

tanks necessary for the propellant as well. The volume necessary for the tanks encompasses nearly

99% of the propellant production plant.

VII. Production of oxygen on the Moon with dynamic solar panels

Oxygen could be produce on the surface of the Moon with dynamic solar panels. The solar flux

is collected by the concentrator, which transfers the concentrated solar radiation to the optical

waveguide transmission line made of optical fibers (Figure 4.). Thus, this high-energy

concentrated solar flux is redirected to the thermal receiver for thermo-chemical processing of

lunar regolith in order to provide oxygen. This system allows a very high efficiency (contrary to

an oxygen production system using electrical heating).

The system consists of three major components:

The concentrator : The concentrator consists of multiple facet parabolic concentrators of

seven 68.6 cm concentrators. The reflectivity of the concentrators over the entire solar

spectra will be 0.9. The concentrators use the secondary reflectors to focus the solar flux

in the optical fiber cables.

The solar power transmission line: transmission efficiency of 0.8. Each transmission line

contained 55 optical fibers made of hard polymer-clad fused silica.

The thermal reactor.

Figure 4.4 Principle of operation of a dynamic solar panel for O2 production

Considering the efficiency above, expected in the next years [1], the all system efficiency is 74

%. The system can achieved a heating up to 2000°C. The system must be equipped with a two

axis solar tracking system.

Primary reflector: 0.7 m

Secondary reflector

Concentrator

Optical waveguide transmission line

Thermal reactor

Regolith Concentrated solar flux: 150 W/cm

2

With an ambient direct solar flux intensity of 880 W/m2, the system with the seven concentrators

give 800 W of power at the output of the optical fibers and 780 W at the quartz output (on the

regolith) which is 150 W/cm2. This flux apply on regolith give a temperature of 2000°C. We

know that 1800 °C, is necessary for the Carbothermalreduction process and melt regolith, thus

this system is efficient enough to produce oxygen.

The primary reflector would be 0.7 m diameter. The weight of one of these systems is 72 kg.

It can produce nearly 98 kg of O2 per year (Table 4.).

Table 4.10 – Sizing of one dynamic solar panel O2 production unit

Moon colony

Concentrator

Number x sizes 7 x 0.69 m

Reflectivity 0.9

Transmission line

Number of fibers 55

Efficiency 0.8

Reactor Heating temperature 2000°C

Whole system

Efficiency 0.74

Power 0.8 kW

O2 (/year) 97.56 kg

Mass 72 kg

The oxygen needed on the Moon is 0.82 kg per crewmember per day. If we consider 9 astronaut

on the Moon colony (3 XM-3).

𝑁𝑢𝑛𝑖𝑡𝑠 =

0.82 ∙ 9 ∙ 365

98= 27.5

(4.1)

To provide enough oxygen for the all colony we need 28 units of dynamic solar powered

oxygen production units. This is a total weight of 1.98 Mg.

References

[1] Takashi Nakamura, Benjamin K. Smith, “Solar Power System for Lunar ISRU Applications “,

Physical Sciences Inc., Pleasanton, CA 94588, 2010