1. 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

2. 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 -

3. 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.

4. 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

5. 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

6. 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

7. 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

8. 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

9. 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

10. 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).

11. 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

12. 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 Carbothermal reduction 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).

13. 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