Up to this day, interplanetary travel is extremely expensive. NASA has a pressing need to lower the cost of interplanetary travel so that it can continue with its interplanetary exploration and research missions. The Moon, our closest neighbor, may provide a solution to alleviate the problem. A Moon station can be an intermediary supply and refill station to such long haul travel. It is desired to design and build a long term Moon station to significantly lower the cost of interplanetary travel. Also, the station may serve to carry research, explore the moon resources, study long term space environmental effects on humans and hardware, and monitor the motion of asteroids in the vicinity of earth orbit.

The primary purpose of this investigation is to design a station on the surface of the moon capable of resupplying, and refueling interplanetary spacecraft in order to reduce the cost of future interplanetary missions, and to extend overall mission lifetime. Secondary to serving as a point for resupply and refueling of interstellar spacecraft, the station would allow for investigating the effect of a low gravity environment on the human body. The station would allow for research into the effects of a low gravity, high radiation environment on organic plant life and technology, and in addition the base could serve as a near Earth object monitoring station. Finally, the station would allow for completion of a detailed geologic survey of the Moon’s surface in order to identify the composition, and any available resources

II. Goals and Objectives

Definition of Objectives

In order to receive a substantial return on investment from the lunar outpost, it will need to fulfill its primary purpose as an interplanetary refueling station, that also has the ability to provide a platform to perform other lunar research for interested international agencies, both commercial and governmental. These research platforms include, but are not limited to, studying the effects of a long term low gravity environment on Earth based life and technology, observing the psychological effects of isolation on base inhabitants, providing a point of view for a near Earth object identification system, harvesting local environment for fuel and other necessary resources, establishing a communication hub for extreme range missions (improve Deep Space Network Communications), and conducting extensive lunar expeditions to survey lunar surface and composition.

The primary function of the lunar outpost focuses on the need for refueling spacecraft for interplanetary missions by harvesting the local environment. Installation, and implementation of this base will assist in reducing the overall cost of future interplanetary missions by providing any necessary fuel requirements for a spacecraft near orbit. Constructing an outpost capable of fulfilling these primary and secondary objectives will call for a set of resources and constraints that need to be met for the project to succeed. The first of these requirements is to identify and call to the principal players that would contribute and benefit from the overall mission.

Principal Players

The lunar outpost will by no means be a small endeavor to construct. Principal players in general are the primary contributors towards the completion and implementation of the mission altogether. The return on investment is the key factor towards the level of involvement of the players for the base itself. Seeking to reduce the cost for future interplanetary missions would draw the attention of the major space agencies around the world. Key agencies considered for the base include both privatized and government funded entities. These agencies include:

National Aeronautics and Space Administration (NASA)
Canadian Space Agency (CSA)
China National Space Administration (CNSA)
European Space Agency (ESA)
Italian Space Agency (ASI)
Indian Space Research Organization (ISRO)
Russian Federal Space Agency (RFSA)

Each organization above has a stake in gaining invaluable research and usability out of the outpost. The United Nations, and United States Congress would have a higher stake in funding and policy reformation regarding the nature of the mission. Privatized agencies such as SpaceX and Blue Origin have much to gain in supplying the mission with the spacecraft capable of transporting the base structure and supplies upon the lunar surface. Success of the mission will hinge on the receptivity of the principal players toward the ultimate goal of interplanetary travel.

Timescale

Defining the mission timescale can be separated into four phases; concept exploration, detailed development, production and deployment, and operations and support. All considerations for this mission timeline are based on the technology available during the initial design process as well as the financing for the overall program. The concept exploration stage will be spent on determination of the preliminary requirements for the base. This includes ideal base location, and base construction, as well as the feasibility to accomplish the primary, secondary and tertiary mission objectives. From these consideration extensive cost analysis and financial planning will need to be taken into account. The project will then be brought to the principal players whose aim it is to benefit from the mission as a whole. The majority of time spent in this phase will be allocated to acquiring funding from the overall process. The estimated duration for this phase will take around one to two years’ time.

The detailed development phase of the project contains the overall system components and technology developments for the base equipment and logistics of the mission. Here the base designs, vehicle payloads, and orbital mechanics will be taken into consideration. This phase of the project will account for the longest duration of the mission timescale with an estimated duration of six to eight years. At the same time, a separate team will begin exploring and conducting the required research to make the project marketable. This will include Quenchgun prototyping, lunar composition mapping, and feasibility of lunar soil fuel synthesis. Production and deployment of the project will commence once all design analyses are completed. Construction and manufacturing of base components, Quenchgun and other assemblies required will last about four to five years in duration.

Pre supply missions will then take place to launch base materials along with the dredger excavator drone to begin initial base excavation. Once initial pre supply launches have begun, the base construction will last for a 2-year duration. By the time the base trenches are fully excavated and set in place, the researchers will be sent to location and stationed on the base, finishing final construction and beginning their initial research and fuel synthesis duties. Once completely constructed and operational, the base is capable of up 20 years of operational use. The following table, i.e. Table 1 describes the mission timescale.

Table 1. Mission Timescale

Phase	End Defined By	Estimated Duration
	Concept Exploration
Concept Planning and Funding	Sufficient funding acquired and general plan laid out	1-2 year
	Detailed Development
Overall Design	Payload and base designs, necessary orbital trajectories established	4-6 years
	Production and Deployment
Production and Manufacturing	Base, Quenchgun, components manufactured	4-5 years
On site excavation and base construction	Pre supply, Drone excavation and base construction commences	2 years
	Operations and Support
Operations	Fuel Synthesis and Resupply, Low G Research, Near Earth Object Monitoring	20 year Operational use

III. Requirements and Constraints

Initial Requirements and Constraints

Concerning initial requirements and constraints that dictate mission success, the outpost must be able to meet the following basic needs to operate in its intended manner:

The outpost must be able to support up to four researchers for up to one year at a time in a safe and secure environment.
The base itself must be able to operate 20 years without requiring any heavy maintenance.
The base construction process must be adaptable to future improvements and base expansions when necessary.
The outpost’s capability of synthesizing and delivering up to at least 12,000 kg of fuel to the lunar orbit each month (provided fuel transportation and synthesis technologies meet the requirements).

The first three requirements listed above call for a need in conducting the structural analysis in order for operation of the base to commence safely and successfully upon the lunar surface. The dependence on the advancement of technology during time of implementation should be of important note when discussing the needs of mission success. Overall mission cost will be the largest constraint that must be met. Initial cost analysis shows the mission total to a factor of $100 Billion USD. Further investigation shows a total cost of up to $182 Billion USD. A detailed cost analysis describing the breakdown is presented in later sections of this study.

Resources

Identification of the resources available in the environment that are accessible and utilizable to the base will determine the operational and functional limits the outpost will face during its 20 year operating life cycle. Regarding fuel synthesis and transportation, the environment around the outpost will need to be studied, and surveyed thoroughly. The composition of the lunar regolith across the moon shows abundance of oxygen and other trace metal elements that can be utilized for fuel synthesis. Frozen hydrogen and even possible water deposits within the many craters across the lunar surface also pose as possibilities of resource mining for fuel synthesis. Even the regolith itself can serve as radiation shielding for the outpost depending on the nature of base construction. These available resources are highly dependent on the location of the lunar outpost, whether it will be stationed along the lunar equator or on either of its poles. This discussion will continue in the alternative concepts later in the present study.

System requirements

In order to ensure the base has operational status for its life cycle, a series of system requirements must be met in terms of communications, power, and structural needs of the outpost. These requirements include, but are not limited to, the following:

5 MB/sec upload/download minimum required for communications
Constant communication with Earth
Power supply for life support, thermal control system (TCS), communications, dredgers, and Quenchgun
Base must be designed to withstand one atmosphere of internal pressure, and maintain structural integrity under the max expected loading with a minimum factor of safety of 2
Base must provide radiation protection up to the standards put forth by the National Council on Radiation Protection & Measurements
Life support to sustain four researchers in comfortable conditions
Micro debris impact protection
Thermal control system (TCS) functionality
Efficiency of supply transportation payloads to the moon

IV. Alternatives

Launch System

The launch system was designed to transport the 450 metric tons of required mass to the lunar surface. Many existing launch vehicles and spacecraft were considered, but the selection was narrowed down two heavier launch vehicles. The SpaceX Falcon Heavy and the Space Launch System by NASA were the two final alternatives. The SLS offers advantages in cargo mass to Trans lunar Injection (TLI) and also has a much larger payload fairing to reduce constraints on cargo volume. The Falcon Heavy has reusability that could help to meet the requirement of lowering overall costs. The Orion Spacecraft and the Dragon V2 were the capsules that were focused on to provide manned transport.

Table 2: Payload to TLI for large Launch Systems

Booster	Payload to TLI
SLS Block 1	25000 kg
SLS Block 1B	~ 36500 kg
SLS Block 2	~ 51000 kg
Falcon Heavy	16000 kg

SLS Payload chart jpeg.jpg

Figure 1: Payload capabilities for SLS configurations [1]

Orbital Trajectory

A Hohmann transfer offers the most fuel efficient transfer for trajectories requiring minimal inclination changes. This would be accomplished by conducting a single tangent burn in LEO putting the spacecraft on a TLI trajectory. Another single tangent burn would be initiated near the moon to enter and circularize a low lunar orbit.

A free return trajectory is a special type of trajectory that is held by a spacecraft to ensure that it can return to Earth in the rare case of propulsive anomalies that could prevent a burn to return to Earth. This trajectory was considered because it can provide another level of safety for astronauts.

Figure 2: Basic TLI 2D orbit mockup

Base Location

Two alternative base locations on the moon that were focused on were Sinus Aestuum Mare and the area surrounding Shackleton Crater. The selection factors for these locations were temperature gradient and daily amounts of solar radiation. Sinus Aestuum Mare is close to the lunar equator reducing delta V requirements. It also has a level, featureless surface. Shackleton Crater located near the lunar South Pole is close to constant sunlight and has a small temperature gradient of around 170K to 260K. A frozen water reservoir in the crater also offers the opportunity for a resource that could be used for fuel.

Final Report.jpg

Figure 3: Moon and its selenological features

Base Structure

Multiple ideas for possible base layouts were considered in initial planning of the project including inflatable structures, underground habitats dug out of the lunar surface, and an orbiting station but were ultimately abandoned due to factors such as cost, safety of the base inhabitants and ease of mission performance. After consideration the best course of action was considered to be a prefabricated structure that would be landed on the surface and buried in lunar regolith using dredger drones that could perform basic construction operations. The drones would excavate a location suitable for landing and then place the final structure in the 4-meter-deep trench. After the base is successfully lowered the regolith would then be placed back on top to act as environmental shielding for the occupants. The alternative structural designs for the base are presented in section VI. Structure.

Fuel Delivery System

Fuel Synthesis

Options for fuel synthesis include water reservoir mining of unknown abundance and oxygen synthesis. Water presence on the moon is volatile, with any liquid or gas quickly becoming. Oxygen is abundant in local regolith and could be used to supply liquid oxygen rocket engines. However, oxygen synthesis is difficult and requires ionization of regolith to separate elements.

Figure 4: Theoretical Quenchgun basic composition [2]

Fuel Delivery

Alternatives for delivery of fuel from the lunar surface to lunar orbit include conventional rocket launches and the use of a Quenchgun. Conventional use of rockets to transport fuel would need to also use the synthesized fuel, but the delta V requirements for the delivery are low. The Quenchgun is energy efficient using only 10 kW of power per 1.5 metric tons of payload. However, the Quenchgun is still conceptual and increases the required mass for the base by around 333 metric tons.

Power Supply

While there exist many possible options for base power supply, the most viable and abundant source of power is from the solar radiation provided by the sun. Ultimately the amount of solar radiation available for the base to utilize depends on the final location. Alternate power supply concepts were considered for this mission, however due to the reliability and abundance of solar energy, the alternative concept for power supply will focus on the available solar cell technology that will be utilized in the entirety of the mission lifecycle.

Figure 5: UTJ Solar Cell

Single Junction Solar Cells

Single Junction Solar Cells can reach an efficiency of 14% because they only have one junction, which means that they can only a small section of the sun’s energy because they only accept a small range of the sun’s wave length. Although they have a low efficiency they are lighter in weight than other solar panels because of this, the density of these cells are 0.81 kg/m2. Although using these lower efficiency cells means that we would need a larger area to acquire the required power, this is not a problem because of the vast area on the lunar surface.

Ultra Triple Junction (UTJ) Cells

UTJ Cells are 28.3% efficient because they have three junctions on top of each other. Each junction can accept a different spectrum of the sun’s wave length, so these cells can accept a higher percent of the sun’s radiation. Because of the increased junction density of the cells, the overall density of the cell at 0.85 kg/m2 is higher than single junction cells. This cells can also last up to 25 years. The efficiency of these cells however outweigh the higher density which is why UTJ cells will be used.

Life Support

In order for the base to be manned by human beings, a life support system is required capable of providing the required food, water, and oxygen needs for daily human metabolic requirements. The system must be capable of long term operation, as a result life support based on resupply methods is not feasible for the present mission. The following right hand figure captioned Resupply vs. Regenerative illustrates this by showing the cumulative mass launched for resupply over time for a sample mission to LEO with four crew members. It is apparent that for a mission with a timeline on the order of multiple years, a bioregenerative closed life support system is best. The different methods for management of gases, consumables, and wastes are illustrated in the following left hand figure above Alternative Life Support Methods, and the way in how these three items are managed will be the defining attribute of the life support system. The present mission intends to base the life support system similar to that of the ISS, but eventually transform it into a Controlled Ecological Life Support System (CELSS).

Figure 6: Life Support Alternatives [3]

Life support will require an initial installation of atmosphere in the base habitat to allow for organic life. The initial constituents of air must be brought from the Earth. Installation of an air filtration system is required to avoid exposing organic life to any toxic gases that might be present in the bases, and to provide a high quality of atmosphere within the base. The levels of carbon dioxide and oxygen must be carefully regulated to allow for sustained life on the base. Ideally, in the long term the organic life would create a bioregenerative cycle in order to eliminate, or at least reduce, the dependence upon Earth for life support related resupply.

Management of consumables and wastes is of concern; a long term option to help with both management of carbon dioxide, and consumables is a lunar farm on the base to produce nutrients required by humans to sustain metabolic activity. The feasibility of any plant based farm on the lunar base is strongly affected by:

Crop light absorption capability
Crop harvest index which is equal to the amount of edible biomass produced divided by total amount of mass of biomass present
Digestible energy per edible unit of biomass
Efficiency in energy conversion

Shown next is the estimated size and energy requirements for a lunar farm provided by the NASA Astrophysics Data System.

Figure 7: Estimated size and energy requirements for a lunar farm from NASA Astrophysics Data System [4]

These requirements allow for an estimate of the feasibility of a lunar farm as a source of food and gas reclamation. The most difficult topic is how to deal with wastes. Ideally 100% recyclability is desired, but is not always achievable. Waste processing is not 100% efficient. Not all human wastes can be used directly by photosynthetic organisms, necessitating waste processing technologies capable of breaking down plant and human wastes. Two options are biological waste treatment, or chemical waste treatment. Presently, biological waste treatment such as is the case on Earth is not possible on the Moon, thus it is more likely a chemical system would be used to break down the wastes molecules into constituents that can then be used for other purposes in the base.

The life support system on the lunar surface must be able to provide 4 lbs/day of food for each crew member, 6 lbs/day of water for each crew member, and 3 lbs/day of oxygen with carbon dioxide being continually filtered from the atmosphere within the base. This can be accomplished using a lunar farm containing organic plant life which could use the carbon dioxide for photosynthesis, and provide food for astronauts. Through water reclamation, and recycling of all materials on site the base can be made a self-contained ecosystem capable for supporting human life. Complete reusability, and recycling of all inputs and outputs to from humans and other organic life in the base system is the primary constraint on a CELSS. Presently, the technology and detailed processes do not exist that are needed to realistically accomplish this method of creating a closed, self-sustaining system. Future research into how this scheme can be made viable will ultimately lead to increasing the feasibility of the proposed lunar refueling, and research station.

Communication

Communications for the Lunar outpost would be derived into 2 sections; Base pre supply and excavation and Habitat communications requirements. During the pre-supply and excavation phase of the mission, existing interplanetary communications such as the ones used on the International Space Station (ISS) would be more than sufficient to send and receive information for not only telemetry and guidance of the incoming spacecraft, but the Dredger drones as well. Direct communication between the ground and the moon would also be available for several hours a day without the aid of a communications satellite in orbit if ISS communications are unavailable.

The system that would likely be chosen is the existing Tracking and Data Relay Satellite (TDRS) system that is currently being used by the ISS. The TDRS uses two separate frequencies to communicate with other orbiting bodies, the S (2-4 GHz) and Ku (12-18 GHz) bands. For data transmission the system uses the Ku band, however for signal acquisition the S band is used due to its larger bandwidth.

After the base is constructed and researchers are actively working on the base, the requirements for communication would be expected to meet the increased needs of a manned station; mission updates, video conferences, and personal messages the base must be able to transmit video in close-to real time, meaning that during peak usage the communications system must be able to handle around 5 MB/s of both download and upload speeds.

Comparative to the ISS, the S and Ku bands will be sufficient enough to establish two-way video and audio communication among the researchers within the base including those in extra vehicle activities (EVA). These bands are also capable of file transfer communication between mission control on Earth or even the ISS if necessary.

V. Structure

In order to assess whether the base structural requirements were met, 3 different structures have been considered. The first structure analyzed is that of the central HAB unit of the base. The second structure represented the cylindrical segments of the base protruding from the central HAB unit, and the third structure analyzed was a segment of the flooring. Following is an early base concept made in SolidWorks showing the external geometry of the base.

Figure 8: Layout of Initial Base Design: External Geometry

Limiting Loads of Interest:

Static:

Body forces resultant lunar gravity
Reaction loading from the weight of the regolith on top of the central HAB
Internal base pressurization
Material creep can alter, and possibly compromise, the tensile properties of the material

Dynamic:

Footsteps of base inhabitants on the flooring can excite the structure
Vibrations from electric power generators can excite the structure
Varying electric currents can cause heat fluctuations within the structure
Fatigue due to cyclic pressure loadings acting upon the structure
Fatigue due to cyclic change in ambient temperature outside of base
Operation of Quenchgun
Operation of central elevator in HAB unit
Payload launch vibrations during initial supply phase

Limiting Loads Analyzed using SolidWorks:

Static:

Weight of regolith atop base
Internal base pressurization

Reasons:

Creep negligible since the operation temperature is less than half the melting temperature, the ambient fluctuations outside of the base are 238.15 K to 253.15 K and the melting temperature of the base wall material is above 855.15 K
Force due to gravity much lower order of magnitude compared to pressure forces

Dynamic:

Cyclic loading and unloading of the flooring to model inhabitant footsteps
Simulation of the acceleration experienced by the solar panels during launch assuming linear dynamics analysis with a base excitation of 5 g’s in the x direction, 0.1 g’s in the y direction, and 0.2 g’s in the z direction. Here x is pointed downward with y and z defined according to the right hand rule and as is shown in the figure
Analysis of the frequencies, and modes of natural vibration of the central HAB unit

Reasons:

Vibrations from operation of power generation equipment and machinery not maintained at resonant frequencies for prolonged periods of time
Launch phase likely represents the highest possible loading conditions of the solar panel, so if it can survive launch it can survive the mission
Assumed relatively small cyclic pressure fluctuations throughout most of base except near airlocks
Fluctuations in ambient temperature are well below half the melting temperature so thermal stresses are negligible
Heat transfer from current carrying wires in base not considered, assumed to be negligible

Simulation Description:

Static Load Analysis:

To ensure the researchers are able to conduct the necessary fuel synthesis for interplanetary refueling in a safe and lasting environment, simulations were used as a preliminary design tool. In order to determine the optimal base model, multiple designs for the central base unit as well as the 4 “HAB extensions” were created and run through multiple loading scenarios. These designs include introducing uniform and nonuniform thickness to the HAB extensions, a honeycomb based structure compared to a constant thickness, a reinforced roof model, and additional rod supports in the model of the central base unit. The HAB extensions will be comprised of 3 layers: an inner Mylar insulated and padded layer, a middle structural layer, and an outer “shielding” layer. The middle layer will be comprised of either the honeycomb or the flat plate structure. The Central HAB unit will only be comprised of the structure itself and a Mylar and padded insulated layer.

Three design options were chosen for the central HAB unit for a static structural analysis. The central HAB unit is the spherical dome structure in the center of the base. Each design utilizes a concentric ring in the roof. The first design was a modification of the original base design, but with a reinforced roof. The other two designs had the same roof design as in the first case, but added reinforcement rods in different orientations.

Initially the central unit was designed without a fortified ceiling. This proved to yield a factor of safety on the order of 1.1. This low margin of safety was not compliant with the standards of required for the lunar outpost, so reinforced rings were designed into the roof to add more structural support. This provided a factor of safety of 1.23 with a max stress of 4.465 *107 N/m2 and a maximum displacement of 9.138 mm. These results were not sufficient to meet the minimal standards of safety.

The primary difference between the three designs is in the supports of the roof. In order to increase the factor of safety, 4 sets of 2 rods were placed along the failure points near the bottom of the doorway of the central unit connected to the reinforcement rings on the roof. This resulted in an increased factor of safety for the entire structure of 3.096, a maximum stress of 1.87*107 N/m2, and a maximum displacement of 0.967 mm. Although this factor of safety was satisfactory and the maximum stress was well below the yielding stress of the material, this design was not the best choice. Despite the increased strength, the rods were not positioned ideally. The rods were placed in pairs at each of the four doorways, and consequently made maneuvering through the doorways difficult.

The final design changed the locations at which the support rods were installed to allow for easier maneuverability for base habitants. The concentric ring roof support was made larger, and the support rods were connected between that ring and floor. This addition resulted in a factor of safety of 2.66, a maximum stress of 2.068*107 N/m2, and a maximum displacement of 1.126 mm. Though the second design provided the most favorable results in terms of failure criteria, consideration of the occupants of the base make the third design the best choice. This design best met both the structural and ergonomic requirements.

Design of the cylindrical sections protruding from the central HAB were analyzed by varying the wall geometry. The two designs exhibited a compromise between using the least weight possible, while staying well below the allowable stress values. Two cases of interest were to use a constant wall thickness, or to use a honeycomb wall structure.

Initially the honeycomb structure was to be designed around the HAB extensions, but difficulty arose in creating a cylindrical geometry with a regular honeycomb wall structure. For simplicity, two flat plates were modeled in place of the cylindrical wall. Each plate was subject to a loading of 100 Newtons. Each flat plate was on the order of 15x20 mm in size. Through use of SolidWorks’ Simulation toolkit, it was determined that the honeycomb structure had a minimum factor of safety of 37.31, a yield strength of 5.515*107 N/m2 , and a maximum displacement of 3.612 microns. The flat plate had much larger margins for safety than that of the honeycomb design. The constant thickness plate provided a minimum factor of safety of 83.41, a maximum stress of 6.611*105 N/m2, and a maximum displacement of only 2.242 microns.

The flat plate provided a much higher factor of safety compared to the honeycomb structure, but for the present investigation a factor of safety of 5 is considered to be larger than is necessary. Hence, the flat plate is not the best choice. The plate weighs roughly 0.2548 kg, whereas the honeycomb design is around 0.0403 kg. The honeycomb design is over 5 times lighter than that of the flat plate. This significant reduction in weight offsets the difference in factor of safety between the two designs since a factor of safety of 5 is satisfactory. Utilization of a honeycomb wall structure would provide a minimal weight while still ensuring the structure is not likely to fail.

The cylindrical sections were also tested using uniform and non-uniform wall thickness to further optimize the structure weight and factor of safety. Using a uniform thickness of 2.5cm and under normal loading conditions the structure had a FOS of 3.3 and max stress of 15.1 kPa. For the non-uniform model, the wall thickness was increased to 10 cm at the sides to decrease the maximum stress, this lead to a FOS of 3.6 and a max stress of 16.4 kPa. Because of the increase in FOS for the non-uniform model it was selected as the design. In combination of the non-uniform design and honeycomb wall structure, the cylindrical section will be able to lower its overall weight while maintaining its structural integrity for the whole mission.

Dynamic Load Analysis:

Structural design requires an investigation into the way a system behaves when subjected to a dynamic loading. It is important to assess the fatigue of the structure over an extended period of time to ensure the desired life cycle can be achieved. Over time the strength of the structure may become compromised resulting in deformation or other unwanted characteristics. The base location is remote, so any necessary maintenance of the base structure must be planned far in advance. Ideally, the base will not need maintenance until the end of its life cycle if designed properly. The mechanical, and thermal dynamic loadings of interest were previously listed under Section II Limiting Loads. Considered in the present study are primarily vibrational excitation due to loading and unloading within the base and vibrational analysis of the solar panel structure during launch. The Lunar outpost must be able to maintain its structural integrity while being subjected to any fatigue or vibrations over the course of its 20-25 year life cycle.

The main challenge with dynamic loading is determining what forces will be applied on the central HAB unit. This area must be able to undergo frequent foot traffic without deforming or buckling even after millions of cycles. In order to test for failure, the maximum weight the floor would experience is multiplied by four in order to simulate the most extreme case. For the force due to the weight of researchers, the mass of each scientist was taken and converted to a force using lunar gravity, resulting in a combined total of 4000 N of force. Even after running the loading one million times the results showed that no metal fatigue would be experienced on the floor or any other vital section of unit. This is because even with the loading of both the pressure and the weight of the astronauts, the stress experienced on the frame of the structure is below that of the stress needed to fatigue Al 6061 to the point of failure. This is justified due to the fact that the maximum Von Mises stress experienced at any point on the section is on the order of 2.07*107 Pascals. The minimum stress needed to deform the structure after 10,000,000 iterations would be on the order of 5* pascals. Due to this information, dynamic studies result in zero fatigue being experienced over operational lifetimes.

One of the greatest sources of vibration and dynamic loading will occur during launch. The solar panels required for the transport spacecraft and the base itself will be required to endure significant acceleration during launch. A linear dynamic analysis was conducted by applying a base excitation of 5 g’s in the downward or X direction with an additional .1 g and .2 g acceleration in the Y and negative Z directions to create an offset in loading. The solar panels were fixed on one end face. Two studies were run with the variation between the two being the presence of a fixed edge on the opposite side of the fixed face. The simulation with the top edge fixed helped reduce displacement. From this result it can be concluded that it will be optimal to fix the solar panels at both ends to another surface in order to help reduce displacements and also stress. The stress calculated was significant but well below the yield strength of the material used, Aluminum 6061. The actual solar panels will contain silicon and other materials that may react much differently but the single material study provides a baseline analysis of what can be expected in the structural frame during launch.

The greatest source of failure, second only to fatigue, is vibrations in a structure. To ensure that the structures would not experience such a failure, studies must be performed for each structure and each variation to determine the natural frequencies. If these natural frequencies match up to normal operating frequencies, then a failure is inevitable and padding or extra reinforcement must be added. The normal operating frequency of our central HAB unit will be anywhere from 0-5 Hz due to the footsteps of inhabitants walking around the interior and from large machinery such as vacuum pumps running only when the situation demands it. From frequency analysis the first mode at which all structures experience resonance appears close to 35 Hz. This value is consistent with all structures and designs meaning that our normal frequency of 0-5 Hz will not cause failure.

The most extreme case of vibrations that the structure will experience will likely come from machinery that runs for short bursts such as a vacuum pump. This operates on the order of 200-300 Hz. As this machinery will be housed in the central HAB unit the frequency studies were expanded to include a larger range of frequencies to encapsulate the vacuum pumps contributions. It can be determined that this frequency should not be maintained for extended periods especially with the non-reinforced design. The second HAB design shows the least amount of modes in this range meaning it is the most stable, but not by a large enough margin to consider it the only option. The chosen design with the struts placed at a greater angle experiences a similar order of modes in the vacuum pumps frequency range, reinforcing the choice of the third design for our HAB structure.

Simulation Results:

Table 3: Summary of Numerical Results from Static Analysis of Central HAB Unit for Internal Pressure of 1 atm and external pressure of 0 atm. Figures of contours of interest follow later.

Static Structural Analysis of Central HAB Unit	Factor of Safety	Maximum Stress (MPa)	Maximum Displacement (mm)	Estimated Manufacturing Cost
Reinforced Roof with No Support Rods	1.23	44.65	9.138	$6,240,000
Reinforced Roof with Vertical Supports Rods	3.096	18.7	0.967	$6,880,000
Reinforced Roof with Angled Supports Rods	2.66	20.68	1.126	$6,670,000

Table 4: Summary of Numerical Results from Static Analysis of Small Rectangular Section of Cylinder Wall subject to a Loading of 100 N

Static Structural Analysis of Flooring	Factor of Safety	Maximum Stress (MPa)	Maximum Displacement (mm)	Mass (kg)
Flat Plate Uniform Thickness	83.41	0.6618	0.0002242	0.2548
Flat Honeycomb Plate	37.31	55.15	3.612	0.0403

Table 5: Summary of Numerical Results from Static Analysis of Entire Section of Cylinder

Static Structural Analysis of Entire Cylinder	Factor of Safety	Maximum Stress (MPa)
Uniform Wall Thickness	3.3	0.0151
Nonuniform Wall Thickness	3.6	0.0164

Table 6: Summary of Numerical Results from Static Analysis of Cylinder

Static Structural Analysis of Cylinder

Factor of Safety

Maximum Stress (MPa)

Maximum Displacement (mm)

Estimated Manufacturing Cost

Uniform Wall Thickness

4.583

12.03

0.3848

$1,820,000

Nonuniform Wall Thickness

3.688

14.73

0.3841

$1,830,000

Figure 9: Von Mises Stress Contours (left column) and Factor of Safety Contours (right column)

Figure 10: Honeycomb vs. Flat Plate structural analysis

Figure 11: Cylindrical Extension. Uniform vs. non uniform walls

Figure 12: Solar Panel Dynamic Structural Analysis, one fixed end

Figure 13: Solar Panel Dynamic Structural Analysis, fixed edge on top

Figure 14: HAB Frequency charts

Design Solution and Implementation

Of the designs tested for possible lunar habitats, case three was selected to be the central habitat and the honeycomb design was selected for the cylindrical extensions. These designs were selected after weight, cost of manufacturing, structural integrity, ergonomic interaction with the stations residents, and effective lifetimes were considered. It is because of these multiple variables that case two was not selected was not selected for the central unit. While it showed a higher factor of safety and a lower maximum stress the cylindrical supports would only provide a meter of clearance at each entrance. This amount of space could prove to be a hazard especially over the long operational lifetime of the mission. For this reason and due to reduced manufacturing costs the third case was chosen.

The chosen HAB design consists of two structural rods placed near the floor of each entrance for a total of eight. These rods run from the floor at an inclined angle away from the entrances up to a circular reinforcement ring running around the roof of the structure. These supports help reinforce failure points located near the entrances and drastically increase the FOS of the entire design.

For the cylindrical sections pressure vessel, the honeycomb case was chosen as a final design. This design would provide support for the atmospheric This is due to the honeycomb patterns mass being a sixth of the solid body estimation it was compared to in figure 9’s flat plate analysis. Despite the 66% drop in the FOS the flat plate analysis showed that it would still be within acceptable tolerances under normal loads. While the computational limitations of this report preclude the possibility of creating a dynamic simulation of a honeycomb design a simple mathematical proof can be shown to explain why it is the superior design based on the geometry of a cylinder.

For the solid plate cylinder of 1 cm thickness the calculated FOS is 4.583. The total cross sectional area of this design is .0939 By doubling the thickness of this design to 2 cm the cross sectional area and by linear relation the mass of the cylinder jumps to .1884, almost twice the area as the 1 cm thick design. Even if the required thickness of the honeycomb design requires twice the thickness to achieve the solid plates 4.583 FOS it will still only take a third of the mass to do it. This reduced mass on four separate parts of the final design greatly reduces the mass that is needed to be launched into lunar orbit and saves millions in launch costs.

The ability to withstand the static and dynamic loading the lunar base will experience in its lifetime is a key factor in choosing a suitable material for the base structure. For the present study, Aluminum 6061 was selected due to its ideal strength to weight ratio. A lightweight material is critical to ensure the lowest total mass of the payload required for transport. This reduces propellant requirements and also keeps the amount of launches to a relatively low number, in turn reducing the overall cost of this mission. Furthermore, Aluminum 6061 is very commonly used on the commercial market meaning its manufacturing practices are well established. Structural characteristics of the aluminum alloy are listed in the table on the next page.

Table 7: Aluminum 6061 characteristic properties

Yield Strength	55,148,500 N/m2
Density	2,700 kg/m3
Fatigue Life	10.57 cycles
Young’s Modulus	6.9 e10 N/m2

The material properties listed in Table 1, specifically the yielding strength, and fatigue life satisfactorily allow for a mission life span of 20 to 25 years, based upon the SolidWorks simulations. Assuming no unexpected loading events stresses the structure, which is not without uncertainty, then mission failure would be unlikely. Though other materials were available for selection, Al 6061 was chosen due to its structural rigidity, or shear modulus, and its cost effectiveness.

While the regolith resting on top of the base will provide the bulk of radiation protection, a thin outer layer of Al 6061 will also be added for supplementary shielding. Lead and other carbon composite based material was considered for the outer radiation shielding for the Lunar base. Lead shielding is extremely heavy to consider transporting which is prohibitive for this mission, while carbon composite based shielding is expensive to manufacture and could possibly be damaged during the long lifetime of the mission. The aluminum layer wouldn’t have been able to provide shielding alone, however using a small layer of Aluminum with regolith atop will provide sufficient protection for the researchers posted on the base. Furthermore, the Aluminum material not only provides structural support, but serves as a radiation shielding mechanism thus making it an ideal choice.

VI. Final Design and Concept

Consideration of the alternatives, requirements, constraints, and cost effectiveness has led to the final design and concept detailed below

Launch System

The SLS Block 2 was selected as the primary launch system because of its payload mass advantage and fairing volume advantage over the Falcon Heavy. The Falcon Heavy could still be used for resupply missions in the future due to its reduced costs from reusability, but it does not offer the best capabilities for the current mission. Using the Block 2 SLS, the cargo payload mass available for launch was calculated to be around 30,000 kg. The Block 2 SLS will lift 51,000 kg of cargo and propellant to translunar injection (TLI). The burn for lunar orbit insertion was calculated to require around 13,034 kg of propellant leaving about 38,000 kg for transfer to lunar orbit. The lander was found to require 7,969 kg of propellant to land the payload. For the ascent stage the lander only needs 66.48 kg of propellant to rendezvous with the next incoming payload. The propellant for both the ascent and descent stages will be provided by the incoming spacecraft. This means that accounting for the total required propellant the cargo payload available for transfer to the lunar surface from Earth will be around 30,000 kg.

The construction and use of a reusable lander was also selected because it will serve to reduce fuel costs. The reusable lander would be aboard the first launch and would be used to land the first payload. After landing the initial payload, the lander would then return to lunar orbit to be refueled by the second payload and then assist it in landing on the surface. The number of times the lander could be reused would likely be limited due to various wear and stress. This limitation must be overcome to a point where the lander could be reused a sufficient amount of times in order to justify associated costs.

The propulsion system for the lander will utilize the Draco series engines by SpaceX as they provide sufficient basic specifications to meet the goals of the design. The SuperDraco is already designed for multiple restarts so it will be able to be modified in order to create a reliable and reusable propulsion system to land payloads. It also has the ability to throttle heavily in order to provide precise control, which is extremely important when conducting the landing. A total of 8 SuperDraco engines will be used in order to provide a highly redundant system that can be used many times. The reaction control system (RCS) will be comprised of 12 smaller Draco engines which will provide small adjustments when necessary to maintain the proper trajectory.

The landing system will be automated in case of communication problems with Earth based ground support. The landing system will still have the ability to be controlled manually from Earth. A one square meter solar array will be sufficient to supply 15-20 kW of power to run propulsion and GNC systems. The dry mass will be kept as low as possible. A 1500 kg dry mass is a target for the design so that less propellant will be required.

A total of 15 number launches will be required in order to transfer the crew and the materials for the base, which was estimated to weigh around 450,000 kg. All cargo will be modified as necessary to fit in the 10m by 31m (1800) payload fairing.

Orbital Trajectory

The Hohmann transfer was elected for the transport of cargo because there were no significant time constraints for the cargo transfer. The manned mission will utilize the free return trajectory to ensure safety in case of a failure. This will help satisfy the most important requirement of astronaut safety.

Table 8: ΔV Requirements

ΔV requirements for Mission
Lower Earth Orbit Injection	7.623 km/s
Hohmann Transfer to Lunar Orbit	4.03 km/s
Lunar Descent	0.53km/s

The four-person crew consisting of researchers will launch in the Orion spacecraft aboard the SLS to Low Earth Orbit (LEO). The second stage will then conduct a burn in order to put the spacecraft on a free return trajectory. Although an immediate return is not planned to occur as this is a one-way transport, a free return trajectory will insure safety for the crew in case of a propulsive malfunction or any other anomalies that would jeopardize a safe and successful journey to the lunar surface.

Other cargo and materials required to construct the base will be launched in a payload fairing atop the SLS to LEO. The second stage will burn to put the payload into a Hohmann transfer with the moon at apoapsis. A free return trajectory is not necessary for the unmanned payload and the time constraint will be less important when considering the transfer due to the absence of life support.

Base Location

Many factors make the Shackleton Crater the better location for the moon base. Shackleton Crater has almost constant sunlight at the rim, which provides benefits for power storage and minimizes temperature fluctuations. The near constant sunlight at the rim and the constant darkness also presents many opportunities to keep materials at their operating temperatures without expending much energy. The opportunity to utilize frozen water in the crater played a big part in the decision to choose the Shackleton crater as well. The biggest issue present with this location is the high inclination a launch vehicle would need, in order to deliver the payload at its final location. This can be mitigated though, because the TLI path can be manipulated to deliver the payload directly to the 90° inclination without a burn at the moon.

Power Supply

Due to the location of Shackleton crater at the lunar south pole, the base will receive nearly constant solar radiation; 80-90% in continuous sunlight per 30-day lunar orbit. The amount of radiation that the area near the crater receives is on the order of 1200-1300 depending on the position of the moon in orbit. The overall power requirement for the entire lunar base takes three major project aspects in consideration; base research and life support equipment, Quenchgun needs, and communications.

The main power drain for the base life support systems would include the power drain for the TCS of the base. The EATCS itself draws approximately 15-20 kW in order to actively regulate the heating within the 103total area of the base. Discussing fuel delivery systems, the Quenchgun could launch a 1.5 tonnes payload to lunar orbit, and complete a launch every 2 hrs using 350 kW of power. This gives the outpost the ability to launch a payload once every 2 ½ days at a power requirement of only 10 kW per launch. Regarding communications, a power budget of 10-20 kW is required for both drone and base communications.

Fuel synthesis from ionizing lunar regolith, and near earth object monitor systems tack on an additional 50 and 10 kW of required power respectively. Overall the main power budget requires a bare minimum of 110-120 kW of power on a daily basis. For redundancy, safety concerns, and preparation for future base development, an additional 20% of the power budget would be required to keep the base within safe operational limits. Tallying up the power budget to a total of 120-140 kW of power required. Described in the alternative concepts portion of this report, the solar panels will be comprised of the C4MJ triple junction Gallium Arsenide cells. These cells operate at a 40% solar efficiency and will last for the full 20 years of the mission life.

Figure 15: Final HAB Designs

lunar base.JPG

VII. Allocation of Requirements to System Elements

Communication

The base will require a communication system that operates on the moon and can provide a upload/download combined bandwidth of 5 MB/s. This will be necessary for communication, constant reporting from monitoring systems, and data upload and download from experiments being performed on the base. This overall bandwidth does not need to be error free, but the system must be capable of an error free bandwidth of 100 kB/s to prevent corruption of expensive scientific data. The moon allows for communication downtime to be very small, and provides the possibility of no downtime from periodic events. The TRDS system utilizing the K and Su bands will fulfill the communication requirements for this project.

Base location

The base requires a location that is moderate in ambient temperature fluctuations, ample

in supply of available sunlight, and easily accessible from orbit. The ample sunlight allows for generation of solar energy needed to sustain the base power systems. The Shackleton crater is the ideal location, potentially meeting all of these requirements.

Base structure

The base must be capable of supporting four human crew members for sustained periods of time. It must provide life support to keep the crew comfortable. The base must also be able to withstand the internal pressurization of the base resulting from the life support system, and also any other applied loadings which are a result of the chosen base design. Due to the loading that would be present on the moon, a base design of aluminum in a honeycomb structure is sufficient to meet the structural requirements while costing much less than the alternatives.

Fuel delivery

The system must be capable of synthesis of fuel on site, and transportation of the fuel to a

lunar orbit for acquisition by a passing spacecraft. The desired amount of fuel transported monthly is to be a minimum of 12,000 kg as listed in the system requirements. To accomplish the fuel delivery, the Quenchgun will be implemented. The power requirement to launch 1.5 metric tons of fuel is estimated to be 10 kW, hence roughly 80 kW is the power requirement to reach the monthly fuel delivery goal.

Launch system

The launch system must be reliable, and have the ability to transport all materials needed

for construction to the site. It also must be able to transport personnel to the site after the initial construction phase has commenced. The launch system for initial supply does not need to be reusable, but after the initial phase every effort should be made to recycle launch systems. This is to keep the base cost effective in order to facilitate the prime objective of long term reduction in overall costs of interplanetary travel. For these reasons the SLS Block 2 will be used for initial launches, a reusable lander will be utilized for transportation to the lunar surface, and the Falcon Heavy will be used for resupply over the lifespan of the base.

Orbital trajectory

The trajectory taken by the launch vehicle from liftoff to orbit about the moon must be

determined a priority to ensure the transport of materials from Earth to the site occurs without any unforeseen difficulties. The associated orbital maneuvers will require fuel in order to provide the necessary delta V to accomplish the transport mission. The delta V and fuel requirements are elaborated upon in the Final Design and Concept section earlier in the present study.

VIII. Cost Analysis

Initial rough estimate of the lunar outpost resulted in an amount just over $100 Billion USD. After extensive cost analysis, the final estimate tallies up to $182.5 Billion. With each phase of the project being extensively researched for safety, along with functional and operational redundancy, the complexities increase the overall cost of the outpost as a whole.

Detailed development contains the overall design and research process of the mission as stated. Research taken place in this phase will focus mainly on lunar fuel synthesis and life support available for the base. As these are 2 main success criteria for the entire mission, financial allocation towards this portion is of utmost importance; $10 Billion is dedicated towards Research. Overall base design, including solar array farm, Quenchgun and dredger has $5 Billion allocated. Production costs for human labor had been estimated assuming 100,000 workers at $100k salary per worker puts the 6-8 years of labor of this phase at $80 Billion. Overall the detailed development phase is $95 Billion.

The production and deployment phase depends on the individual cost of base needs, as well as overall labor and launch costs during the 7-year time period. Using the same assumption as in the development phase for labor, 4 years of labor accounts to $40 Billion. An estimated 15 launches will be required for pre supply and base construction missions, costing $11 Billion overall for launches at Cape Canaveral, Florida. Base materials, Quenchgun, Dredger, Solar panel and communications construction account for $6.1 Billion of phase cost. Total estimated phase cost for production and development adds up to $57.1 Billion.

Overall the entire cost of the lunar research base is estimated to $178.46 Billion, only a 21% increase compared to ISS total cost. For a 20-year operational period this estimate is feasible over an extended period of time. Figures in the Appendix shows tables of both the launch weight breakdown and the cost breakdown of the entire mission. $30.4 Billion in recurring costs, resupply launches per yearly basis, and $148.06 Billion in non-recurring costs, research, production, et cetera.

Table 9: Cost Analysis

Cost analyiss.PNG

IX. Conclusion

The overall mission will by no means be a small endeavor to construct. Success of the project will hinge on the receptivity of the principal players toward the ultimate goal of interplanetary travel. To keep public reception in check throughout the design process, the cost of the base and its usefulness was closely tied to those same parameters for the ISS. The presence of a manned lunar base presents opportunity for meeting goals secondary to the refueling of spacecraft seeking to go further than the Moon. The lunar conditions make research studies regarding the effects of gravitational forces different than Earth’s on Earth based life very appealing. In today’s society, too often short term goals are being prioritized above long term goals. The intrinsic human desire for exploration will hopefully unite the people of Earth under a common vision. That vision, to grow as a species intellectually, technologically, and culturally. Ultimately to explore the cosmos, to delve into the deep unknown vastness of space, and to boldly go where no human has gone before.

X. References

[1] SLS payload Capabilities [Web] Availible: https://directory.eoportal.org/web/eoportal/satellite-missions/o/orion-em-1

[2] NASA (1990) A Superconducting Quenchgun for Delivering Lunar Derived Oxygen to Lunar Orbit [web] Available: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19900012490.pdf

[3] Alferov A.V, Grigoriev A.I, Orlov O.I. Space Medicine (2011, June 1) [online] Available:

http://www.unoosa.org/pdf/pres/copuos2011/tech-07.pdf