Inspiration Mars International Student Design Competition

Two-Person Mars Fly-by Mission

Ryan Gilligan (TL) | Nicholas Filipkowski | Mansoor Mustafa | Gregory Van Zant |

Jacob Whiteman | Blaine Zaffos

The Ohio State University

03/15/14

Astronautical

Reconnaissance

Expedition

Spacecraft to

Mars

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Table of Contents

I. Introduction …………………………………………………………………….……………..5

A. Concept of Operation and Team Member Roles……………………………………...5

B. Mass Budget …………………………………………………………………...……...5

C. Cost Analysis …………………………………………………………………...…….6

D. Schedule ……………………………………………………………………………....8

E. Requirements …………………………………………….……………………..…….9

II. Orbital Analysis ………………………………………………….………………..……..….11

A. Trajectory Design ………………………………………….……………………..….11

B. Re-Entry ………………………………………………………………..……………13

C. Launch Vehicle ……………………………………………….…..……………..…..15

D. Heat Transfer Analysis ………………………………………………………..…….16

III. Booster Interface Adapter ……………………………………………...……………………16

IV. Spacecraft Breakdown ……………………………………………………..………..………18

A. Layout and Sizing …………………………………………………..…………….…18

B. Materials and Structure …………………………………………...…………………20

V. Propulsion …………………………………………………………………….…………..…22

A. Layout and Detail ……………………………………………………..…..…………22

B. Risks …………………………………………………………………………………25

C. Cost ………………………………………………………………………………….25

VI. Power System …………………………………………………………………………..……26

A. Budget and Storage ……………………………………………………..……...……26

B. Solar Arrays …………………………………………………………………..…..…27

C. Wiring …………………………………………………………………….…………28

D. Heat Dissipation ……………………………………………………………………..29

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VII. Attitude Determination and Control …………………………………………...……………30

A. Sensors and Actuators ……………………………………………………………….30

B. Docking Maneuver …………………………………………………………………..31

VIII. Communication …………………………………………………………………….….…….32

A. Deep Space Network ……………………………………………….………….…….32

B. Hybrid Inflatable Antenna (HIA) …………………………………………...……….33

C. Transmitters …………………………………………….………………..…….……34

D. Receivers …………………………………………………………………………….35

E. Risks ………………………………………………………………...……………….35

IX. Life Support ……………………………………………………………………..…….…….36

A. Oxygen Generation ………………………………………………………...…….….36

B. Water Regeneration System ……………………………………………….…..…….38

C. Food Supply …………………………………………………………………...…….39

D. Crew Health and Safety ………………………………………………………….….41

E. Spacecraft Atmosphere …………………………………………………………..….43

X. Risk Management ………………………………………………………………………..….44

XI. References: ……………………………………………………………………………..……45

List of Tables

Table 1: Mass Budget ……………………………………………………………………………6

Table 2: Initial cost estimate using USCM8 model………………………………………………7

Table 3: Data from STK Astrogator for asymptotic target vector of departure and arrival……..12

Table 4: Spacecraft mission departure and arrival information………………………………….12

Table 5: Thicknesses of Various Components of the ARES-M Skin………………………...….21

Table 6: Thruster Performance Specs……………………………………………………………22

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Table 7: Propulsion Layout Summary…………………………………………………………...25

Table 8: Propulsion Cost Estimate……………………………………………………………….26

Table 9: Power Budget…………………………………………………………………………...27

Table 10: Solar Power System Characteristics…………………………………………………..27

Table 11: DSN Aperture Fee Tool calculation for DSN use in $2018FY……………………….33

Table 12: Differences in Cost for DSN Payment Methods………………………………………33

Table 13: Additional Needs for Oxygen Generation Process……………………………………38

List of Figures

Figure 1: ARES-M Proposed Schedule …………………………………………………………8

Figure 2: Plot showing Earth-Mars free-return opportunities……………………………………11

Figure 3: Arrival conditions for a perfect launch and delay of 10 minutes……………………...12

Figure 4: Spacecraft trajectory for entire mission………………………………………………..13

Figure 5: Profile of “G-Force” vs Altitude during Re-entry……………………………………..14

Figure 6: Mass vs. Energy MATLAB fit for Falcon Heavy……………………………………..15

Figure 7: Booster Adapter………………………………………………………………………..17

Figure 8: Booster Adapter Specifications………………………………………………………..17

Figure 9: Internal Spacecraft Layout…………………………………………………………….18

Figure 10: External View…………………………………………..…………………………….19

Figure 11: Structural layout of the Apollo spacecraft's skin…………………………………….20

Figure 12: Observation of the inner and outer shells of ARES-M………………………………20

Figure 13: Busek Hall Effect Thruster…..……………………………………………….………23

Figure 14: Fuel Tank and Support Structure…………………………………………………….24

Figure 15: Propulsion System and Plumbing Layout…………………………………………....24

Figure 16: Solar Panel…..………………………………………………………………………..27

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Figure 17: Solar Panel Deployment..…………………………………………………………….28

Figure 18: Wiring Scheme……………………………………………………………………….28

Figure 19: Wiring Channel Separator……………………………………………………………29

Figure 20: Mercury Capsule Retro-rocket Assembly……………………………………………32

Figure 21: Placement and Orientation of the Hybrid Inflatable Antenna………………………..34

Figure 22: Hybrid Inflatable Antenna, µTx-300 Ka-Band Transmitter and NEC Ka/Ku-band

RCVR's…………………………………………………………………………………………..35

Figure 23: “Life Box” illustrating the daily needs of an average human being…………………36

Figure 24: Inflatable Aeroponic System designed by NASA……………………………………40

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I. Introduction

A. Concept of Operations

The Aeronautical Reconnaissance Expedition Spacecraft to Mars (ARES-M) team is designing a

two person fly-by mission to Mars in 2018 as safely, cheaply, and simply as possible. The main

goal of the mission is to demonstrate that humans can not only survive but remain healthy during

extended duration (>1.5 years) deep space missions while also demonstrating the ability for

humans to successfully travel to Mars and back. Remaining healthy includes both physical

health, which considers effects of living long term in a zero gravity environment as well as

mental health, which considers extended duration in a confined living space and isolation from

society. A successful Mars flyby requires an orbit trajectory that will take the spacecraft to Mars

and back to Earth in a reasonable amount of time as well as a propulsion system and launch

vehicle that can provide the energy to make such an orbit possible. An attitude determination and

control system is essential to keep the spacecraft on the correct trajectory. During the trajectory

the astronauts will need sufficient food, water, and clean air as well as access to exercise

equipment to remain healthy. Other components of keeping astronauts healthy will be

maintaining a reasonable temperature of the living space and protecting the astronauts from solar

radiation. The spacecraft will have to be able to communicate with Earth throughout the mission

for a variety of reasons including receiving assistance in the event of an emergency. The various

systems on the spacecraft will require a supply of electricity and therefore an adequate power

source. All these capabilities will be required throughout the more than 18-month mission.

Team Roles

Ryan Gilligan works as the project manager, propulsion system designer, and heat transfer

analyst. Nick Filipkowski works on the communications system, serves as structural designer as

well as conducting thermal analyses. Mansoor Mustafa is the life support system and reentry

analysis specialist. Greg Van Zant works as the attitude determination and control engineer and

designed the experiments to accomplish throughout the mission. Jacob Whiteman is the orbit

determination and modeling expert and responsible for choosing a suitable launch vehicle.

Blaine Zaffos works as the structural and power systems specialist. Together they are going

where no man has gone before, the Red Planet.

B. Mass Budget

After analyzing launch vehicle and thruster capabilities, a total mass budget of 20,000 kg was

decided upon. This mass allows the mission to be completed using one launch which eliminates

many complications associated with multiple launches, such as an orbit assembly of the

spacecraft or having to wait at least a month between launches to prepare the launch pad. It also

allows for the launch vehicle to contribute to the escape energy required for insertion to the

trajectory to Mars. On a side note, the absolute limit of mass the Falcon Heavy can insert into

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orbit is 53,000 kg; however for a Mars journey this number falls to only 13,000 kg. As a result,

the spacecraft requires its own propulsion system that will provide the additional thrust required

to reach Mars. The energy contributed by the launch vehicle significantly reduces the propellant

mass requirements of the spacecraft’s propulsion system. A summary of spacecraft component

masses is presented in Table 1.

Table 1: Mass budget

System Mass (kg)

Power 756

Propulsion 2692

ADAC 340

Life Support 3000

Communications 100

Structure and Other 13000

Total Mass 19889

C. Cost Analysis

The cost estimate below is merely a rough order of magnitude estimate on how much the ARES-

M will cost to create. Most of the estimates are based on masses of different elements; however,

the Flight Software estimate was modeled after the Apollo mission specifications and results

since the ARES spacecraft has similar features. [1] Since Apollo had a restraint on SLOC count

due to hardware requirements at the time, they defaulted to analog inputs and multiple teams of

operators. For this two-person mission the SLOC count will be increased by a factor of 100 in

order to make it operable by the two-person crew. Lastly, the Falcon Heavy Launch vehicle cost

estimate was found directly from the SpaceX data sheets.

The cost analysis diagrammed on the following page was created using the USCM8 model found

in the textbook “Space Mission Engineering: The New SMAD.” [2]

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Table 2: Initial cost estimate using USCM8 model

The cost is also likely to be higher than predicted as a manned space mission incurs extra cost

and complexity due the increased risk of keeping the passengers safe throughout the mission.

Cost increases as more testing and documentation is required for almost all of the spacecraft

components to ensure human safety. The increased complexity and mass is largely driven by the

inclusion of a life support system. The life support includes the food, water, medical supplies,

exercise equipment, and environmental control and regulation that the crew needs to not only

survive but also remain healthy.

The expected cost for the life support system was based on an AMCM (Advanced Mission Cost

Model) created by NASA. [3] Factoring in the weight of the subsystem, its complexity, the

mission type, and its newness, an expression was formed that estimated the anticipated cost:

with each variable assigning a numerical value representing one of the aforementioned

characteristics. This value was adjusted for inflation and the combined sum of all subsystems

was categorized under the life support section.

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D. Schedule

Figure 1: ARES-M Proposed Schedule

Above is the proposed schedule for the ARES-M spacecraft, starting from the early design

phases and ending at the launch date. The first phase is the introduction of the mission statement

and defining the main objectives and purpose of the endeavor. Design concepts and ideas are

discussed amongst team members and trade studies are conducted. From what the mission

requires and the concepts presented during this phase, a basic project approach is constructed and

possible avenues for progression are laid out.

The second phase shifts towards the actual preliminary design of the spacecraft, covering vital

components and major subsystems. Greater detail regarding concepts and design features is

achieved and the skeleton of a basic model that meets the necessary requirements is created. New

technologies discussed in the first phase are researched and possible solutions are found. The

third phase brings together all the major ideas presented in the second phase and refines them to

a higher order. Further emphasis is placed on detail and the overlying concepts for all systems

are completed (both major and minor). Thorough analytical analyses of all spacecraft systems are

performed, including efficiency studies, structural analyses, survivability tests/scenarios, etc. The

final design phase is completed when a project that is ready for initial fabrication is reached.

The fourth phase starts with the fabrication of subsystems and test components. Through

experimentation, studies are performed that validate the effectiveness and feasibility of

components as part of a whole. As each aspect of the design is tested and approved, each part

begins its integration within the entire assembly and an end product is achieved. This goes

through a final verification test to ensure all mission requirements can be fulfilled with the

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completed design. Preparation for the launch begins shortly after as the fifth and final phase is

reached. The launch date for the ARES-M is January 5, 2018.

Throughout the course of the project lifecycle, critical design meetings and reviews are

scheduled that assess the progress. These serve as checkpoints to determine whether or not the

project as a whole is moving at the right pace and if completion will occur as expected.

E. Requirements

The only requirements provided by Inspiration Mars and Mars Society are as follows: “The

requirement is to design a two-person Mars flyby mission for 2018 as cheaply, safely and simply

as possible. All other design variables are open... You are free to select from any technology,

launch vehicle, or flight system that is currently operational or which can be plausibly argued to

be potentially operational by 2018.” Additional requirements to ensure a successful mission are

defined as follows:

Payload and launcher must be at final operational capability by 5 January 2018.

Spacecraft must be autonomously guided.

Mission must be completed using only 1 launch.

Propulsion System

o Propulsion system must provide enough thrust to propel craft through free return

trajectory.

o Propulsion system must have at least 50% redundancy in case of engine failure.

Life Support System

o Oxygen must be generated to fulfill daily needs for a crew of two.

o Water must be recycled continuously to ensure constant supply is available.

o The cropping system must produce the required food to fulfill nutritional needs of

crew.

o Internal piping systems must function properly to monitor the cabin atmosphere.

o Sufficient radiation protection to counter possible dangers in space flight.

Orbital Mechanics

o Spacecraft must successfully leave the Earth sphere of influence and travel to

Mars.

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o Entire orbit must be under 3 years to minimize radiation exposure and life support

elements.

Structure and Materials

o Choose advantageous layout and materials for the spacecraft’s body.

o Structure of spacecraft must have adequate thickness for insulation and load

bearing.

o Plate spacecraft with reflective material to assist with thermal control.

Communications

o Communication system must provide a maximum data rate of >50 Mbps.

o Fit spacecraft with antenna with Ka-Band communication capabilities, preferable

deployable

o Transmitters and receivers must be capable of sending and receiving the desired

Ka-Band frequencies

Power

o Power system must supply a continuous source of electric power to the spacecraft

for the mission duration.

o Power system must contribute and distribute electrical power in correct voltages

to necessary spacecraft components.

o Power system must support the power requirements for average and peak

electrical loads.

ADAC

o ADAC system must keep spacecraft in desired orientation for long duration

engine burn and maximum solar array and communication antenna capability

throughout mission.

o ADAC system must allow for astronauts to manually control spacecraft in case of

emergency.

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II. Orbital Analysis

A. Trajectory Design

After initial concept discussion about propulsion and orbital trajectory, it was decided that any

significant burns when maneuvering around Mars would push the mass beyond an acceptable

limit due to excessive fuel necessary for such burns. Due to the strict limit on mass to complete

the single launch requirement, the mission became a Mars free-return mission. Free-return, as the

name implies, is a specific type of Mars flyby mission in which the initial momentum and

gravity assist from Mars are enough to safely get the spacecraft to Mars and then propel it back

to Earth, with propulsive forces only needed for small course corrections. During the trajectory

the spacecraft will fly within 200km above the Martian surface with around 10 hours in the

100,000 km range in which the astronauts can see Mars.

Patel et. al. determined various times when mars free return trajectories are available and these

are plotted in Figure 2 [4]. Free returns with a time of flight under 1.5 years only occur twice

every 15 years and the one ARES-M will be utilizing is circled in red. This leads to a very tight

schedule and deadlines, and creates a very real risk of the Falcon Heavy not performing up to the

standards assumed here for this mission.

Figure 2: Plot showing Earth-Mars free-return opportunities, Patel at al [4]. The opportunity being used here is circled in red.

The STK Astrogator tool by AGI was used to calculate and model the entire spacecraft

trajectory. In the interest of producing the most accurate solution, the final trajectory was run

using a fully heliocentric model and considering the gravitational effects of all planets and the

sun at all times. Once an initial targeting sequence had gotten the spacecraft close to Mars, the

trajectory was forced to target the B-Plane of Mars and then that of Earth for the return.

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Table 3: Data from STK Astrogator for asymptotic target vector of departure and arrival

Table 4: Spacecraft mission departure and arrival information

The table above shows the exact dates of departure, flyby and arrival. In fact these dates are quite

exact, almost down to the minute. For example, if the spacecraft were to launch just 10 minutes

late, it would end up five million km from Earth, without the added use of trajectory correction

by the attitude thrusters. Although the final result with the added use of attitude correction

thrusters would be much less severe, the added fuel needed for this extra burn time would

severely impact the mass budget regardless. This is another reason the launch date cannot be

changed and the schedule is very tight.

Figure 3: Arrival conditions for a perfect launch (left) and delay of 10 minutes (right)

In addition to the tight schedule due to sensitivity of launch parameters, the other major concern

when discussing manned missions far from Earth is the possibility of an abort protocol. In this

case, abort would need to be very soon in order to have enough time to reverse thrust and come

back to Earth. Due to the long burn times required from the low thrust of electric propulsion

engines, it would take a significant amount of time to not only counteract the thrust from the

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launch vehicle, but then gain enough thrust to propel the spacecraft back to Earth. For this

reason, it is safe to assume that once the spacecraft has burnt through around 10% of its fuel, a

return trip will full burn in the opposite direction would not give enough momentum to make it

back to Earth, in which case it would be better off just continuing on the mission. However, more

analysis with STK is necessary to determine exactly how long into the mission the spacecraft

must abort in order to return safely to Earth, if an abort is even possible.

Figure 4: Spacecraft trajectory for entire mission (green)

Shown above is the ARES-M trajectory in relation to the orbits of Mars, Earth and Venus. The

spacecraft leaves Earth and proceeds in a counterclockwise direction, with a flyby of Mars, and

then returns back to Earth. The Venus orbit is shown simply to verify that the planet itself is not

near the spacecraft during the brief time it passes by Venus’s orbit.

B. Re-Entry

In regards to the return capture at Earth, the optimal reentry speed is to be set at 13 km/s based

on the tests done on the new NASA Pica heat shield. [5] According to STK calculations, the

ARES-M will approach the Earth’s atmosphere with 13 km/s relative velocity, so reentry should

be practical. However should the spacecraft enter at a higher velocity, an aero-capture maneuver

will be used to slow it down until the necessary speed is achieved. Fortunately, this maneuver

will not need to reduce the speed significantly, therefore, it will not increase the overall mission

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time appreciably. In terms of re-entry position and altitude, any altitude below the threshold of

56km will result in a direct descent into Earth atmosphere, as anything above that will initiate the

aero-capture maneuver discussed earlier. According to heat shield calculations discussed later, an

optimal reentry angle of 1.5 degrees will keep the shield from burning up long enough to

transport the crew back to the Earths surface.

With an approach velocity of 13 km/s at an angle of 1.5°, the descent into the Earth’s atmosphere

and a final landing will take approximately 12 hours to complete. The extremely shallow angle

was chosen to minimize the immense loads the re-entry vehicle would experience with such a

high incoming velocity. Slightly raising or lower the flight path angle had a significant impact on

the load profile, emphasizing the sensitivity of the parameter. Applying a simple time step

approach and calculating the drag on the return module during descent, the following plot was

constructed to illustrate a basic load profile:

0 20 40 60 80 100 1200

1

2

3

4

5

6

7

8

9

10

Altitude (km)

G

Figure 5: Profile of "G-force" vs. Altitude during re-entry

The maximum expected “g-force” during re-entry is about 9.6 g’s. Although slightly high

relative to previous manned space missions, the value still falls below the maximum allowable

limit for humans (about 12 g’s). Approximating the critical buckling load of the module using a

truncated conical assumption and comparing to the maximum anticipated loads, a margin of

safety greater than three orders of magnitude was found. Increasing the flight path angle above

the nominal 1.5° would consequently result in greater loads, comprising both the safety of the

crew as well as the structural limits of the return module.

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Thermal loading is another aspect of re-entry that requires proficient analysis to ensure the safety

of the crew members. Using a similar approach to the drag calculations, a thermal loading profile

during descent was created and comparison to the limitations of the Dragon capsule was

completed. Referencing information regarding the thermal properties of the PICA material used

for the heat shield on the capsule, an ablation rate was found with respect to the amount of

thermal load present (W/m2). [6] For the expected loads during the descent of the return module,

a minimum thickness of 1.5 cm would be required to counter the heat. The Dragon capsule is

equipped with a heat shield thickness of approximately 7.6 cm, allowing for a suitable safety

margin. This is needed to ensure the glue and other materials between the crew module and the

shield itself remain cool enough to stay intact throughout the descent.

Upon completing the full descent trajectory, the module will land in the Pacific Ocean where

Naval vessels will be awaiting their return. The crew members will be safely rescued from the

capsule and returned back to the U.S.

C. Launch Vehicle

The state-of-the-art Falcon Heavy from SpaceX will be utilized as the launch vehicle, as it was

the most feasibly launch ready system that was able to lift the payload to the required orbit.

However, since the launcher itself is not scheduled to be tested for a few more months, data has

to be extrapolated from the SpaceX website, which is based on previous launches using the same

Merlin engines.

The analysis of this Falcon Heavy launcher begins with the pure amount of energy (C3) that

would be needed to correctly place the spacecraft into its trajectory to Mars. According to the

STK Astrogator values determined above the energy required for insertion to the Mars-bound

orbit is approximately 38.8 km2/sec

2. However due to the newness of this technology, a plot of

available escape energy the Falcon Heavy could produce for various payload masses was not

available from SpaceX, and was calculated purely by utilizing a spline fit in MATLAB code with

the three points takes from the SpaceX data sheet on the Falcon Heavy.

Figure 6: Mass vs. Energy MATLAB fit for Falcon Heavy

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It can clearly be seen from the plot above that when the launcher is given the payload mass of

this mission (22000 kg) the energy is not even positive, meaning the spacecraft will not even

leave the escape orbit. Through calculations done with STK Astrogator it has been calculated

that an extra ΔV of approximately 3.5 km/s is needed from the spacecraft’s propulsion system to

insert itself into the correct trajectory.

D. Heat Transfer Analysis

In order to determine that the cabin of the spacecraft could be maintained at 293 K for astronaut

survival and comfort, a steady state heat transfer analysis was performed on the spacecraft at 2

different positions in the orbit, its closest and furthest points from the sun. The outside

temperature of the spacecraft, which is lined with a few atom thick layer of gold due to its

desirable absorptive and emissive properties, was calculated using T=(GsαF12/εσ)1/4

where Gs is

the solar flux (W/m2) at the respective location, α is the absorptivity of the material, F12 is the

view factor between the sun and the spacecraft, ε is the emissivity of gold, and σ is the

Boltzmann constant [7]. Then the heat transfer between the inside of the spacecraft and its

outside were determined by considering conduction through the spacecraft structure and

radiation through its protective layers of multi-layer insulation (MLI). A thermal circuit was used

to determine the heat transfer between the inside of the craft and the outside surroundings. The

spacecraft is closest to the sun when it passes through part of Venus’s orbit in its trajectory. The

spacecraft is furthest from the sun when it is at Mars. Spacecraft insulation was sufficient at both

points in the trajectory and the spacecraft will have to provide or reject no more than 5 Watts of

heat at either point. The spacecraft will reject heat at Venus by providing air conditioning at mars

and reject the heat from the air conditioner using thermal straps and a radiator. The craft will

provide heat at Mars using space heaters. The MLI is very effective at preventing heat transfer

via radiation through the craft and make maintaining a desirable temperature for the astronauts

well within the capabilities the life support system.

III. Booster Interface Adapter

The main requirement for the booster adapter is to safely support the spacecraft during launch as

well as being as light in weight as possible. The booster adapter is inspired by the booster adapter

design for the Orion spacecraft which holds similar requirements. The preliminary material being

used will be aluminum in order to maintain minimal weight. Further analysis will be completed

to see if the structure and/or the material will be suitable for the mission. The adapter will need

to adapt the 3 meter diameter spacecraft to the 5 meter diameter fairing in order to hold the

spacecraft. The structure will be primarily composed of cylindrical beams, receiving the majority

of their stress in axial loads. Below is the general design of the adapter.

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Figure 7: Booster Adapter

Figure 8: Booster Adapter Specifications

Figure 8 provides a more detailed drawing of the adapter as well as the size and weight

specifications.

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IV. Spacecraft Breakdown

A. Layout and Sizing

Figure 9: Internal Spacecraft Layout

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Figure 10: External View

Figure 9 and Figure 10 show the interior and exterior layout of the spacecraft, respectively. The

extreme rear of the spacecraft contains the main propulsion boosters and tanks. This section

contains possible hazards, and was placed at the end to limit any damaging effects on the crew or

any other part of the craft. This is followed by the general storage area where water, oxygen,

nitrogen, and hydrogen are housed. Note that the water storage has an enclosure within the tank

to allow crew members to protect themselves in the emergency of a solar flare. The main life

support machines are placed in the next section, including the Water Recovery System (WRS)

and the Oxygen Regeneration system. The urinal is attached to the WRS to facilitate the

transport of urine to the processing unit; the toilet is nearby for similar reasons. A small hand-

wash showering station follows within the “bathroom” area and is in close proximity to the WRS

to allow for wastewater recycling. On the opposite side, a small area for food preparation/storage

is giving which will contain necessary dining items and potable water for meals. This is

positioned near the main living area containing the crew sleeping quarters as well as an

entertainment section with a TV for video messages and viewing media. The open space in the

middle has the potential to encompass seating arrangements or setups for other leisure activities.

The living area is followed by the exercise section containing a treadmill, bike, and resistance

machine. Crew members will spend a large part of their day in this part of the spacecraft so

ample space was allotted for this. The main food supply section appears near the top of the

layout with plenty of room for the Inflatable Aeroponic Systems designed by NASA. The crew

will be tending and harvesting crops daily from this region so enough space was given to

accommodate the expected traffic. The extreme front of the craft contains a SpaceX Dragon

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Capsule which will primarily be used for re-entry back to Earth. It contains many of the flight

and control instrumentation and acts like a command section.

The exterior of the spacecraft contains four large solar panels designed to meet the power

requirements of the mission. Small control boosters are seen on the outer layer to account for

minor course corrections and spacecraft orientation.

B. Materials and Structure

Structural Layout

The structure of the spacecraft is based on the layout of the Apollo vehicle. As shown in Figure

11, the outer skin of the Apollo involves a honeycomb layer of stainless steel, sandwiched

between two sheets of the same material. The inner skin follows the same pattern, except an

aluminum alloy is used. In between the two shells is a layer of insulation [8].

Figure 11: Structural layout of the Apollo spacecraft's skin[3]

The outer shell of ARES-M will have a circular cross section, while the inner shell has an

octagonal cross section. This was done so that equipment on the inside of the spacecraft could be

placed on flat surfaces. Figure 12 presents a visual of the model.

Figure 12: Observation of the inner and outer shells of ARES-M

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Materials

Instead of stainless steel and aluminum alloys used by Apollo, the ARES-M will be constructed

using a graphite-epoxy composite material. This material was chosen due to its high strength and

low weight, especially relative to steel and aluminum. The insulation will be multi-layer

insulation (MLI) which will help protect the crew from radiation in addition to insulating the

living module.

Due to the fact that the spacecraft will be oriented with one side of the spacecraft always exposed

to the side, the exterior of the spacecraft will be plated with polished aluminum. This material

was chosen in place of gold due to the controversial and difficult nature of plating composites

with gold. According to Table 22-10 in “Space Mission Engineering: The New SMAD”, for a

cylinder with insulated ends at Earth, with a solar constant of 1366 W/m2, the exterior

temperature will be approximately 23oC. Realistically, the spacecraft’s exterior temperature will

be higher than 23oC, but the high reflectivity of the polished aluminum will help to keep the

exterior temperature as close to the ideal case as possible. The exterior of the spacecraft will also

have a series of radiators to keep the mechanical components, such as the wiring and fuel tanks,

at their optimal temperature. The layout of the radiators will resemble the layout of the Apollo

Service Module.

Thickness

The thickness of the inner shell is 0.02 meters (0.787 in.) and the thickness of the outer shell is

0.031 meters (1.213 in.). Due to the octagonal shape of the inner shell, the overall thickness of

the two shells and insulation varies. At points where there is minimal insulation, at the “corners”

of the inner shell, the insulation thickness is 0.0127 meters (0.5 in.). At points where there is

maximum separation between the inner and outer shell, midway between corners of the

octagonal shell, the insulation has a thickness of 0.0254 meters (1 in.). This results in a range of

overall thickness from 0.0637 meters (2.5 in.) to 0.0764 meters (3.0 in.). Table 5 summarizes this

information.

Table 5: Thicknesses of Various Components of the ARES-M Skin

Structure Thickness (m) Thickness (in)

Cylindrical Outer Shell 0.02 0.787

Octagonal Inner Shell 0.031 1.213

Insulation (Minimum) 0.0127 0.5

Insulation (Maximum) 0.0254 1

Overall Thickness Range 0.0637 – 0.0764 2.5 – 3.0

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Risks

The skin of the ARES-M is thinner than the Apollo spacecraft that it is modeled after. However,

ARES-M uses stronger, lighter materials. It is expected that use of graphite-epoxy composites

will provide enough strength to make up for the lower thickness, but only further analysis can

prove this assumption. If the strength of the spacecraft’s skin is found to be inadequate,

additional thickness will be added to the shells as necessary.

V. Propulsion

A. Layout and Detail

Given the large delta v requirements for interplanetary travel, high specific impulse thrusters

became desirable to limit the amount of fuel necessary. Liquid thrusters have benefits such as

high thrust and low burn time but after analysis would have required a prohibitive amount of

propellant at their current operational capabilities. Thus electrical propulsion became the only

design option.

After performing trade analysis for various electric propulsion systems, eight Busek 8 kW Hall

Effect thrusters were chosen for the mission. Performance characteristics for the BHT 8000 are

shown in Table 6. Burn time was estimated using the ideal rocket equation: tb=m*V/F. Where

tb is burn time, m is spacecraft dry mass, and F is thrust. Eight thrusters firing produces 4.056N

of thrust which requires a burn time of 111 days; the solar panels can provide 70 kW of power up

to 128 days into the mission. If four engines fail, the spacecraft can still make it to Mars with a

221 day burn time. These results have also been verified by utilizing the high fidelity modelling

of STK and the corresponding finite burn propagation elements. The finer details of the program

itself are a bit too complicated at this time; however it was shown clearly that the long burn time

does not affect the trajectory in an unrecoverable way.

Table 6: Thruster Performance Specs. Ref Busek.com

BHT 8000

Thrust 507 mN at 8 kW

Propellant Xenon

Isp 1880 s

Mass 20 kg

Size 24x24x10 cm

23

Figure 13. Busek BHT 8000 Hall Effect Thruster. Ref http://busek.com/index_htm_files/70008511B.pdf

The ideal rocket equation used the spacecraft dry mass, orbit V, and thruster Isp to predict 2218

kg of Xenon propellant for the mission. A nominal tank pressure of 3000 psi was used in the

Ideal Gas law, PV=mRT, to determine the volume requirements and thus the .78 meter radius of

the circular fuel tank. The gas constant was for the Ideal Gas equation was determined from the

universal gas constant and molar mass properties of Xenon, the mass was determined from the

ideal rocket equation, and the temperature was assumed to be at the internal cabin temperature of

293 K. Figures 14 and 15 show the mounting design of tank. The spherical tank will have a

circumferential skirt of triangular tabs that will rest in the Aluminum support ring shown in

Figure 14. The tank is made of a thin .5 cm layer of Aluminum 7075, which has a density of

2800 kg/m3 and maximum tensile yield stress of 368 MPa [9]. A 1.65 cm layer of a carbon fiber

composite, which has a density of 1628 kg/m3 and maximum tensile yield strength of 474 MPa,

is wrapped around the aluminum [5]. Using the lightweight but strong composite allows for the

minimization of tank mass. Tank pressure and geometry were used to calculate the minimum

thickness of the tank using the following equation: where p is tank

pressure, r is radius, and σmax is the maximum tensile yield strength of the tank material [9]. The

tank support stand and the tank inside its support stand are illustrated in Figure 14. Note half the

composite is cut out to display the inner aluminum shell of the tank.

24

Figure 14: Fuel Tank and Support Structure.

Figure 15: Propulsion System and Plumbing Layout.

The plumbing of the system is defined by eight ¼ inch stainless steel tubes leading from the tank

to the engines. The mass flow rate is low enough to where the pressure drop throughout the lines

is negligible. Solenoid valves and flow regulators will be used to deliver the correct mass flow

rate to the engines. Since different engines may be firing at different times each line will also

contain isolation valves to separate the tank and thruster outlet. A summary of propulsion

requirements and component masses is provided in Table 7.

25

Table 7: Propulsion Layout Summary.

Number of Active Thrusters 8

Delta V 1.94 km/s

Burn Time 110.72 days

Propellant Xenon

S/C Mass w/o Propellant 20000.00 kg

Propellant Mass 2218.44 kg

Thruster mass 160.00 kg

Fuel Tank mass 313.36 kg

Total Propulsion Mass 2691.8 kg

B. Risks

If more than 4 thrusters fail, the spacecraft will not have enough thrust to complete its free return

trajectory. Also, if any valves fail, then the mass flow rate will be incorrect resulting in change in

thrust and engine performance. In order to mitigate these risks the propulsion system is a

completely redundant system. The spacecraft has enough power and thrust with 4 engines to

complete the trajectory to Mars and back; however using 8 thrusters reduces the burn time while

also leaving redundancy in the system in the event of engine failure.

If the Xenon pressure vessel bursts, then it will result in catastrophic system failure due to

release of all available propellant. To mitigate this risk, high strength but lightweight Al 7075

and a carbon fiber composite was chosen for the tank material. Also, an additional 20% thickness

was added to minimum thickness required by the max tensile yield stress of the materials.

C. Cost

Estimating the cost of the propulsion system was difficult because we could not receive quotes

from manufacturers. However, a conservative estimate based on online research of costs is

presented in Table 8 [10].

26

Table 8. Propulsion System Cost Estimate.

Component Estimated Cost

Xenon Cost $2,218,435.56

Fuel Tank and Support Structure $50,000

8 BHT 8000 H.E. Thrusters $ 12,000,000.00

Total $14,268,435.56

VI. Power System

A. Budget and Storage

The spacecraft power system will not only be used to power all onboard systems but it will

also be responsible for powering the electric thrusters. The power system is designed to hold

a 10% contingency as a safety feature to ensure the spacecraft has the proper amount of

power throughout the mission. The manned mission systems, life support, and

communications will require about 6 kW of power. The propulsion system and contingency

will require the rest of the available power which will minimally be 32 kW for the propulsion

system for a total of 41.8 kW to power the spacecraft.

As the spacecraft travels away from the sun, the solar load decreases inversely by the radius

from the sun squared measured in astronomical units (au). Earth is located at 1au from the

Sun and the maximum distance from the Sun during the Mars flyby will be 1.386 AU. This

number is significantly smaller than it could be, all the way up to 2au, because Mars will be

at Perihelion at the time of the mission. Due to the decreased solar load, the solar arrays will

need to be larger than if used only in Earth orbit to provide the necessary power when at

Mars. The size will be based off of the power required multiplied by 1.92 (1.3862). Because

of this, there will be an excess of available power early in the mission when the spacecraft is

closer to the sun allowing the power system to power more electric thrusters and therefore

reducing the total burn time.

Solar arrays will be used due to the high cost and limited availability of RTG’s. RTG’s can

also provide a radiation hazard to the astronauts onboard, cause protests from environmental

advocates, and are generally avoided for launch systems. To completely power the spacecraft

when at Mars, the arrays will be designed to hold 80.3 kW of power when at Earth. Excess

Power will be stored on Lithium Polymer batteries as they are the most efficient and lightest

in weight.

27

Solar Array Area (m^2) 210.2032272

Solar Array Mass (kg) 588.5690362

Battery Mass (kg) 167.2

Spacecraft System Power (kW)

Manned mission systems/life support 5.25

Communications and Control 0.75

Propulsion 32

Contingency (10%) 3.8

POWER REQUIRED 38

POWER AVAILABLE 41.8

POWER BUDGET

The spacecraft will be equipped with four planar Gallium Arsenide Ultra Multi-Junction

(GaAs Ultra MJ) solar arrays. These particular solar panels produce approximately 382 W/m2

and weigh 2.8 kg/m2. Tabulated below in Table 9 is the power budget as well as the solar

array specifications.

Table 9: Power Budget

Table 10: Solar Power System Characteristics

B. Solar Arrays

The solar arrays will be folded when stored for takeoff and will deploy when in space

travel. They will be deployed in an accordion fashion similar to a scissor lift powered by an

electric motor. Depicted below is the solar array as well as how it will appear when

deploying. Each array will be approximately 3x18 m in size.

Figure 16: Solar Panel

28

Figure 17: Solar Panel Deployment

As stated above, Lithium Polymer batteries will be used due to their efficiency and light

weight. A drawback to Lithium Polymer’s is the chance of battery combustion. This issue is

mitigated by placing the batteries in different locations around the aircraft to isolate a

potential issue if it occurs while minimizing damage to other batteries. Lithium Polymer’s

achieve approximately 250 Whr/kg resulting in a total mass of batteries, including a 10%

contingency, of 167.2 kg which is depicted above in Table 10.

C. Wiring

The wiring for the spacecraft will need to be as simple as possible to reduce the possibility of

failure as well as make it as easy to repair as possible. The main scheme will include having

channels running down each side of the spacecraft. These channels will be categorized for

wires depending what they are for and what they are running to. The channels will be

accessible behind removable panels in the case of malfunction.

Figure 18: Wiring Scheme

29

Above in Figure 18 is the concept for the wiring. The red strips represent the channels running

through the spacecraft, while the green line represents the wiring from an individual component

to the channel. Below in Figure 19 is what the channel separator will look like. The cut out

semicircular shapes on the sides will be where wires are inserted into the channel as well as

outgoing. The ribbed section on the inside will separate the wires for simplicity. This way,

individual components or multiple similar components will be in the same category.

Figure 19: Wiring Channel Separator

D. Heat Dissipation

All of the electronic equipment will require heat dissipation to maintain a habitable environment.

An equation representing the amount of heat that will need to be dissipated is [7]

Gs is a solar flux constant approximated to 1418 W/m

2, σ is the Boltzmann Constant 5.67x10

-8

W/m2*K

4, absorptivity α is approximated to 0.4 (white surface), and emissivity ε is estimated as

0.8. Qw is the amount of heat that will need to be dissipated, estimated at 1kW. To maintain a

0.5m2 area for the radiator the temperature will be 487K.

30

VII. Attitude Determination and Control

A. Sensors and Actuators

With the current configuration of the spacecraft the Moment of Inertia matrix, in units of

kilograms per meters squared, is:

Cubes of varying densities were placed inside of the spacecraft model to represent the interior

components, such as the fuel tank.

Since the entire spacecraft consists of two separate modules, two sets of attitude determination

and control systems are required. Both the crew module and the living module will be three axis

stabilized. It is very important for the attitude sensors and actuators in this mission to be very

accurate, less than 1 degree, because the flyby orbit to get to Mars for this mission has little room

for error since it will have to return to Earth. The main pointing concerns for this mission are for

the solar panels, to allow for maximum power, communication antenna and, most importantly,

spacecraft pointing while during the long duration burn from Earth to Mars. Because of these

three factors, the spacecraft will have to be oriented in such a way that one entire side of the

spacecraft will always be facing towards the Sun. Because of the large surface area exposed to

the Sun, there will be solar radiation torque on the spacecraft. To calculate the solar torque

applied to the spacecraft at Earth and Mars the following equation was used,

(1)

where is the solar constant, 1366 W/m2 at Earth and 588 W/m

2 at Mars, c is the speed of light,

is the sunlit area, q is the reflectance factor, taken to be 0.8, is the center of solar pressure,

is the center of mass, with taken as 0.2 meters, and is the incidence angle of the

Sun. The solar torque at Earth was calculated to be 4.05 x 10-4

N-m and 1.05 x 10-4

N-m at Mars.

Since a majority of the time will be spent in interplanetary space, solar radiation pressure is the

only environmental torque that is taken into consideration.

The crew module, which will house the astronauts during liftoff and reentry, will have a zero

momentum control system. The system will consist of three rate gyroscopes and twelve reaction

control thrusters. The three rate gyroscopes are required to manage all three axes. The twelve

reaction control thrusters will be laid out in the same way as the Apollo Command Module for

31

complete control of roll, pitch and yaw. The main purpose of the crew module’s control system

is for minor course corrections just before reentry after the crew module has been undocked from

the living module. Cold gas thrusters will be used for the reaction control thrusters because of

their simplicity and reliability. Higher impulse fuels, specifically monopropellant hydrazine and

liquid oxygen liquid hydrogen bipropellant, were the initial choices but were decided against

because of their added risk. The crew module thrusters won’t be fully utilized until the final days

of the mission, which means the fuel will have to be safely stored for 500 days. Using cold gas

thrusters is a much safer fuel choice.

The living module, where the crew will spend a majority of their time during the mission, will

provide the attitude control throughout the mission until it is jettisoned from the crew module

before reentry. The attitude determination sensors chosen are three rate gyroscopes, one for each

axis, six sun sensors, two for each axis to cover both positive and negative, and one star tracker.

The rate gyroscopes and sun sensors will be used together in the beginning of the mission to

initially stabilize the spacecraft, and then the star tracker will be used to achieve higher accuracy

attitude control. The actuators for this mission will be sixteen pulsed plasma thrusters (PPTs) and

four control moment gyroscopes (CMGs). The four control moment gyroscopes will be placed in

the rear of the living module, in the same area as the Xenon tank for the main propulsion system.

The control moment gyroscopes are essential to providing continuous and highly accurate

attitude control during flight. The pulsed plasma thrusters will be arranged in clusters of four

with each cluster placed at 90 degrees on the exterior of the spacecraft to provide full attitude

control capability. The main purpose of the pulsed plasma thrusters will be for CMG

desaturation procedures. Pulsed plasma thrusters were chosen for this mission instead of

chemical propellant thrusters because the spacecraft’s power output capability is already large

because of the use of Hall Effect thrusters for the main propulsion system. When the PPTs need

to be used for CMG desaturation, power can be routed from the Hall Effect thrusters to power

the PPTs. Pulsed plasma thrusters have much higher specific impulse, approximately 1500

seconds, than their chemical thruster counterparts, usually between 200 to 400 seconds, and do

not require large propellant tanks, which saves weight.

B. Docking Maneuver

Due to the configuration of the vehicle at launch, for the crew to move from the capsule into the

living module, the capsule will have to undock from the living module, turn 180 degrees and re-

dock with the living module. In addition to its twelve reaction control thrusters, the crew capsule

will have a strap on retrorocket assembly, much like the Mercury capsule, to provide the forward

thrust for this maneuver. Figure 20 shows the assembly on the bottom of the Mercury capsule,

covering the heat shield.

32

Figure 20: Mercury capsule retrorocket assembly [19]

To approximate the delta V and propellant mass required for this maneuver, the Clohessy-

Wiltshire equations and the ideal rocket equation were used. Assuming that the spacecraft are in

a 700 kilometer orbit around the Earth, the time to complete the maneuver is 20 minutes, and the

maximum separation between the two vehicles is 100 meters, the delta V required is

approximately 0.1808 m/s. Having hydrazine as the fuel for the rocket assembly, with an Isp of

235 seconds, the propellant mass required for this maneuver is approximately 0.3298 kilograms

of hydrazine. After the docking maneuver is complete the rocket assembly can be jettisoned from

the capsule.

VIII. Communication

A. Deep Space Network

ARES-M will be utilizing the Deep Space Network for Ka-Band frequency communication

between the spacecraft and earth. This system has been proven to meet the requirement of >50

Mbps based on historical data [11]. NASA's Jet Propulsion Laboratory presented the DSN

Aperture Fee Tool to calculate the cost of using this system, taking such factors as antenna use,

tracks per week, and time of flight into account. A screenshot of the evaluation software is

presented below in Table 11, showing the calculation of the 2018 Fiscal-Year lump sum cost

method.

The communication link on the trek to Mars is expected to have 80.92% visibility, while the

available link on the return to earth is expected to have 99.72% visibility. In addition, the 70

meter antennas of the DSN will only be utilized 25% of the time for high data rate

communication, including video feeds and conference calls. The remaining 75% will use the 34

meter antennas for low data rate communication, such as basic signals and commands. These

percentages represent the most cost efficient method for use of the DSN as presented in

Krikorian et al [11].

33

Table 11: DSN Aperture Fee Tool calculation for DSN use in $2018FY [12]

Since the duration of the flight traverses over 2018 into 2019, the cost for this procedure has a

slight variation. Payment can either be calculated as a lump sum paid during the fiscal year of the

mission’s start or paid over the fiscal years encompassing the mission’s entire duration, in this

case 2018 and 2019. This fact results in two different payment options: a lump sum paid in 2018

or payment spread over 2018 and 2019. Obviously, the cost differs with the change of the fiscal

year. This is illustrated in Table 12 below.

Table 12: Differences in Cost for DSN Payment Methods

Payment Method Cost

Lump Sum $2018FY $85,465,994

Payment Spread over $2018FY-$2019FY $86,212,273

Payment of all DSN costs in a lump sum at the beginning of the mission will cost $746,279 less

than spreading out the cost over the duration of the mission. For this reason, the lump sum

method will be utilized. Although it may be advantageous to spread payments of the Deep Space

Network over a longer period of time, the money saved provides the largest benefit to the

mission. Since this mission is expected to have a very high cost, any budget trimming is

encouraged.

B. Hybrid Inflatable Antenna (HIA)

The Hybrid Inflatable Antenna (HIA) will be used aboard the ARES-M. This 2-meter antenna

will be stowed during launch and deployed once the spacecraft is in orbit, which eliminates

concerns over payload size and launch loads on structurally weaker parts of the antenna. The

34

antenna is currently approved for X-Band frequency transmissions, but will meet the

requirements for Ka-Band transmissions by 2018 through planned modifications [13].

There will be two antennas included aboard this mission for a couple of reasons. The first is

redundancy, in case one of the antennas malfunctions during deployment. The second reason is

to ensure that there is always at least one antenna facing earth. The antennas will be mounted

near the middle of the spacecraft, 4.47 meters upwards from the base. They are also placed in

perpendicular alignment with the solar panels to avoid interference. They are attached to a

foldable, movable arm which will help orientate the antennas to constantly face earth. The

placement and orientation of the anterior antenna is shown in Figure 21.

An exact cost was unable to be found for this item. Therefore, this part will be given a value that

will be expected to cover its cost, $10 million. A lower number was chosen in part due to the

HIA's mission of giving a low cost option for spaceflight communication. Since there are two

antennas, this will result in a total cost of $20 million.

Figure 21: Placement and Orientation of the Hybrid Inflatable Antenna

C. Transmitters

The transmitters that will be stored on ARES-M are µTx-300 Ka-Band Transmitters, provided by

Space Micro Incorporated. There will be two placed on board for redundancy purposes.

Together, they have a mass of 2.72 kg, with each having a power requirement of +28Vdc±6

Volts.

35

Each transmitter costs $1.2 million, in order to meet the requirements for a NASA Level 1 part

assurance [14].

D. Receivers

The receivers used on ARES-M are Ka/Ku-band RCVR's provided by NEC. There will be two

placed on board for redundancy purposes. Together, they weigh 1.86 kg, with each having a

power requirement of 8.9 Watts nominally in steady state and 10.4 Watts nominally in transient

state [15]. In order to attain an estimate for this product, it was assumed that the receiver would

be slightly higher priced than the previously mentioned transmitters. The rationale is that they

are both communication equipment for deep space travel, however, NEC is a Japanese company

as opposed to the American Space Micro. Therefore, it is expected that shipping, exchange rates

and foreign labor would increase the price. It will therefore be assumed that a Ka/Ku-band

RCVR will cost $1.4 million. There will be two Ka/Ku-band RCVR's for redundancy purposes,

resulting in a final cost of $2.8 million.

Figure 22: Hybrid Inflatable Antenna, µTx-300 Ka-Band Transmitter and NEC Ka/Ku-band RCVR's [13], [15], [16]

E. Risks

Risks of the Deep Space Network on earth are minimal, since there are multiple antennas at

multiple locations. The system has also been in use for many years and has been proven as a

reliable system.

Although it has been stated that the HIA will be ready for Ka-Band frequency by 2018, this still

presents a risk. Should the antennas not be ready by the launch date, a different antenna will be

required or different transmitters and receivers for X-Band communication will need to be used.

Use of X-Band, however, would not be ideal since Ka-Band is better suited for high data rate

communication [13].

There is a risk that a transmitter or receiver will malfunction. However, the inclusion of a second

transmitter and receiver is intended to alleviate this risk. Use of a transmitter that meets NASA

Level 1 part assurance requirements also helps alleviate this risk by confirming a transmitter of

the highest possible quality.

36

IX. Life Support

The life support system required for crew sustainability can be split into the following categories:

oxygen generation, water regeneration, food supply, crew health, and spacecraft atmosphere.

Subsystems for each section are designed to accommodate a crew of two with technology

available by the 2018 launch date. An overview of each subsystem addresses the feasibility of its

inclusion on a manned mission to Mars. The figure below (“Life Box”) illustrates the average

daily intakes of a human being and serves as a benchmark for various calculations and design

choices mentioned in upcoming sections. [17]

Figure 23: “Life Box” illustrating the daily needs of an average human being

A. Oxygen Generation

From the Life Box figure, an average human being requires approximately 0.83 kg of oxygen per

day for normal operation. For a two person mission, the requirements are doubled to 1.66 kg.

The oxygen generation method for this mission will be composed partly of systems currently in

use in the ISS as well as advanced systems still in the developmental phase. Combining these

systems creates an overall process composed of a four phase closed-loop cycle. [18]

The first step in the cycle is splitting water using an electrolysis process. This essentially takes

one molecule of water and separates it into individual components of hydrogen and oxygen via

37

an electric current. This process is shown in (2) found below. An example with 12 molecules of

water is shown to match values found in later expressions. In these chemical reactions, the ratio

of input and output elements is more important than the number of molecules themselves. With

the molar mass of one water molecule being 18 g/mol while hydrogen and oxygen being 1 g/mol

and 16 g/mol, respectively, the expression in (3) was formed:

(2)

(5)

As seen in (4), an initial value of 2 kg of water was chosen as it produced the ideal amount of

oxygen, seen in (5). This process will require the oxygen generation machine currently used on

the ISS to perform the electrolysis phase. Even with a 95% efficiency for this system, enough

oxygen is still generated to adequately supply a crew of two (1.66 kg required). The excess

hydrogen from this reaction will be stored in tanks for use in later phases.

The second step in the cycle will be using this created oxygen as part of the human respiration

cycle. In this process, an equal amount (in terms of molecules) of oxygen will be respired and

carbon dioxide exhaled, as demonstrated in (6) for the purely gaseous form of the reaction:

(6)

An inner piping system within the spacecraft will be required to collect the carbon dioxide

present in the air. A system that filters the ambient air and separates it into necessary components

will need to work in tandem with this air extraction piping network. This concept will also be

referenced in latter sections discussing strategies for water regeneration.

The third step will combine hydrogen and carbon dioxide, the two unused parts of the previous

steps, using a Sabatier reaction. This process will mix the two byproducts and generate methane

and water as an output, as seen in (7). However, the amount of hydrogen required for this

reaction is twice as much as the hydrogen generated in the first step. As a result, the initial cycle

for this oxygen generation method will require additional hydrogen to complete this phase.

However, later iterations of this process will have the necessary hydrogen through the fourth

phase (discussed later).

38

Twelve molecules of water are produced from the Sabatier reaction, matching the requirements

for the electrolysis step. A system that effectively performs this process is still being tested by

NASA but should be suitable for use within a few years. This will work in conjunction with

electrolysis machine to directly transfer the excess water to the subsystem.

The excess methane from the Sabatier reaction will undergo pyrolysis, a method that performs

chemical decomposition using heat. This will place the methane in extreme temperatures and

separate the compound into carbon and hydrogen components, seen in (8) below:

The pyrolysis reaction will require a chamber with an operating temperature range of 1000°-

1200°C to perform. The excess solid carbon will be removed from the spacecraft and the

produced hydrogen will be supplied to the Sabatier reaction. This process was tested at NASA’s

Jet Propulsion Laboratory and operated at an impressive 95% efficiency.

Inefficiencies are expected from any process and will therefore necessitate additional water to

counter losses within the overall process. With 95% efficiency rates for electrolysis and

pyrolysis, as well as an assumed efficiency of 95% for the Sabatier, the following values were

computed:

Table 13: Additional Needs for Oxygen Generation Process

Required Produced Daily Needs Mission Needs

Water 2.000 kg 1.805 kg 0.195 kg 97.5 kg

Hydrogen 0.422 kg 0.402 kg 0.020 kg 10.0 kg

In addition to these added masses, the weight of the subsystems and tanks will contribute to the

overall mass of the oxygen generation system.

B. Water Regeneration System

The average daily requirement for a human in regards to water consumption is approximately

3.53 kg, as seen from the “Life Box” figure. This total includes water from normal drinking

needs as well as from food sources. Due to the long duration of the mission, it would be

infeasible to carry all the necessary water on board the spacecraft for a crew of two. This would

simply create additional mass constraints that would further compromise the delicate mass

balance for the mission. As a result, a water regeneration system will be implemented to recycle

the used water and create a near continuous cycle.

The research facilities at NASA Ames are currently developing a highly efficient Alternate

Water Processor. [19] Rather than the previous design that required urine distillation and a series

of filtration steps to produce potable water, this alternate system relies heavily on forward

39

osmosis to complete the task. This uses a semi-permeable membrane to allow water to pass

through while blocking larger molecules (waste) from proceeding, effectively separating the

components. The liquid is drawn using an osmosis gradient that creates a low and high

concentration solution to generate flow. This new system, still in the developmental phase for

space flight, is anticipated to have a recovery rate of over 95% according to engineers. Older

systems used both a separate Urine Processing Assembly and a water recovery system, the

former having an efficiency of about 70%. This alternate has a clear advantage in its overall

water recovery and provides a suitable option for an extended mission in space. [20]

Water will be supplied to the processor from a few sources, the main supply coming from urine

and wastewater transported from the crew urinal/toilet system. Wastewater from hygiene and

other uses will also be gathered and supplied to the main system for processing. In addition,

humidity from the ambient air will be extracted and cycled through this machine and recovered

as potable water. Water lost from respiration and sweat will be gathered and recycled from this

process. This subsystem will be associated with the main environment control system that

maintains suitable living conditions for the crew. Air will be passed through pipes across the

spacecraft and a key task will be filtering the air into useful components (carbon dioxide for

oxygen generation, humidity for water regeneration, balancing air composition, etc.). The

combination of these various sources will generate the majority of daily water needs for the crew

members via the Alternate Water Processor.

C. Food Supply

For a two person mission to Mars with an expected duration of 500 days, carrying all the

necessary food will add significant weight to the overall mass budget. To reduce the weight

footprint from food, an Aeroponic cropping system will be implemented and operated by the

crew to produce edible crops throughout the mission. Aeroponics, similar to hydroponics, does

not require a medium such as soil for the growth of plants. Instead, the roots of the plants are

suspended from the body within a chamber and frequently sprayed with a nutrient solution in

misting intervals. This supplies the nutrients directly to the plant and results in faster and

healthier growth. Compared to hydroponics, an Aeroponic system requires 98% less water since

the plants are not actually submerged in the water-nutrient solution. These characteristics offer

Aeroponics a clear advantage over conventional practices and make it a viable option for the

mission.

NASA researched the growth of adzuki beans using an Aeroponic cropping systems in space in

the late 20th

century. They determined that this method produced crops more rapidly and yielded

a healthier plant, potentially offering more nutritional value as well. The system was said to be

simple in its operation as the drawbacks of a soil based medium from growth were not present

(contaminants, pesticides, etc.); therefore no extensive/expert knowledge would be required to

maintain an Aeroponic system. NASA recently engineered an Inflatable Aeroponic System (AIS)

40

that offers a lightweight, inflatable structure to house the cropping system. [21] This is would be

well suited for a long duration mission as mass trimming is of high concern.

As mentioned earlier, an Aeroponic system relies on interval misting to supply nutrients to the

crops. Using trends found in common Aeroponic systems, a misting frequency of 10 second

bursts in 10 minute intervals is suitable for healthy growth. [22] The chamber comes equipped

with a recycling system that transports unused solution back to the main misting apparatus. For a

35L/h spray rate, water consumption seems extremely high in the initial stages. However, plants

generally transpire 95% of absorbed water during biological processes, meaning a near majority

of the water will be recycled for future use. From the misting rate and recovery options for

unused/transpired water, an estimated 350 kg of water will be required to operate the Aeroponic

system for a 500 day mission. A small portion of the mass will be allotted to the nutrient mixture

since only a minute concentration is present in the overall solution.

Analyzing the daily needs for an average human, a cropping scheme including lettuce, rice, and

beans will be implemented to provide crew members with necessary nutrition. Lettuce will

supply essential vitamins and fiber to their diet, rice for carbohydrates, and beans for protein.

Additional nutritional needs will be satisfied via small portions of pre-packed food. An offset

cropping method will be used to grow plants separated in one week intervals for a four week

total period. This would ensure that crops are ready to harvest weekly by the crew and a

continuous supply is present for the future. Since all nutritional needs will not be met through

crops, additional food will be stored to reinforce necessary requirements. However, this should

still be a limited amount as a large portion of daily needs is addressed through the Aeroponic

system.

Figure 24: Inflatable Aeroponic System designed by NASA

41

D. Crew Health and Safety

Crew health and wellness can be split into both a physical and mental component, as the strains

of a long mission in space create issues in both aspects. In addition, safety precautions for

radiation fluxes like solar flares have to be taken to protect crew members from potential risks.

From a physical standpoint, the effects of muscle degradation are heightened in a zero gravity

environment as bone and muscles become weaker. Crew members need to be physically fit to

perform the necessary daily operations required during space flight. Failure to counter these

effects could have life-long impacts on the health of astronauts exposed to prolonged missions.

Personnel at NASA recommend at least 2.5 hours of daily exercise to reverse muscle atrophy;

even more might be needed for a longer mission. To accommodate these requirements, the

spacecraft will house common exercise equipment currently in use at the ISS. This includes the

Cycle Ergometer, Treadmill, and Resistance Exercise Device. The Cycle Ergometer is similar to

an exercise bike with the ability to measure vitals for members to monitor health. The treadmill

is of typical form with additional harnesses to prevent participants from floating off. The

Resistance Exercise Device is a weight lifting machine that offers a wide range of exercises for

several key muscle groups, tailored with straps for a zero gravity environment. A combination of

these machines should provide the crew with the essential daily exercise needed to maintain a

desired level of physical fitness.

Since the overall mass budget of this mission does not allow for any heavy scientific equipment,

the experiments will have to use existing equipment that will already be found in the living

module, such the exercise equipment and laptops. The proposed Mars fly by mission will be the

longest duration mission for a single crew in the history of spaceflight, he current record is 438

days, held by cosmonaut Valery Polyakov aboard Mir, and will be the first time humans leave

the influence of the Earth-Moon system. Therefore, the experiments to be conducted during the

flight will focus on the physical and mental toll such a long spaceflight will have on the human

body. On the physical side, the research that is being done on the International Space Station

concerning the effects that long duration spaceflight has on the human body can conducted

during the fly by mission. Using the exercise equipment aboard the spacecraft, the astronauts will

be monitored to help doctors and trainers create the most efficient workout regimen for long

duration spaceflight. Since the goal is for humans to eventually walk on Mars, a workout routine

that will help minimize the effect of zero gravity during flight so the stress and strain of a

reintroduction to gravity on the surface of Mars won’t affect the astronauts as greatly.

Another major area of research will be studying the mental state of the astronauts and how they

react to being in a confined space for 500 days. When on long duration spaceflights on space

stations astronauts are able to be reminded of home just by looking out the window down at

Earth. This will not be the case for the two astronauts on the fly by mission. Only candidates that

prove that they are strong mentally, as well as physically, will be selected for the mission, but the

42

long term isolation and realization of the scale of this mission may have a profound effect on the

crew. Close observation and study of the crew by psychologists will provide security for the

well-being of the crew and provide extremely valuable information on how humans can cope

with long duration space travel.

From a mental standpoint, being separated from family and home (Earth) for 500 days can create

possible homesickness problems. To counter these effects, the spacecraft will be designed with

the necessary equipment to relay pre-recorded video messages from family and friends. Seeing

and hearing loved ones from back home will provide crucial moral support and motivation to

crew members on their journey. In addition, designing the interior of the spacecraft to model a

“home” setting will reduce the impression of flying in a metal cabin for a year and a half. This

can be achieved by simply decorating the interior with items reminiscent of home.

Doctors will determine what medical supplies are feasible and necessary for the mission. One of

the astronauts will preferably be a medical doctor in the event of a medical emergency. However,

if this is not possible there will be doctors on call that can instruct the astronauts what to do via

the communications system in the event of a medical emergency. Though this may not be an

ideal situation the astronauts must understand and accept the risk of being millions of miles from

the nearest hospital.

A major concern with any manned space flight is the risk of radiation. The possibility of solar

flares and galactic cosmic rays present a real danger to the health of crew members as both

contain high energy particles. Based on the requirements set forth in the mission statement and

space radiation cancer risk studies done by NASA, a 500 day deep space mission to Mars would

result in approximately a 5% increase in the probability of excess cancer. [23] However, this

value strictly assumes nominal shielding from the aluminum structure of the spacecraft. For the

ARES-M spacecraft, multi-layered insulation used for thermal control also serves as a great

radiation shield. Based on a radiation shielding analysis done by NASA, a GCR exposure for a

500 day period with MLI shielding would yield an approximate fatal cancer probability increase

of 1.5%, assuming 60 rem equates to a 1% increase in risk. Compared to the expected dosage

from an unshielded analysis, the MLI shielding itself reduces the overall radiation within the

spacecraft by about 70%. [24]

To record the radiation activities within the spacecraft, monitoring systems will be placed

internally to measure the radiation levels. Externally, radiation measurement techniques used by

the ISS such as Neutron Detectors and Directional Spectrometers will provide insight on the

magnitude and direction on incoming rays and allow crew members to better prepare for the

situation. [25] Studies show that consuming antioxidants after radiation exposure help reduce the

dangers that follow. Fortunately, lettuce and beans (both grown on board) are great sources of

antioxidants and should assist in countering the radiation effects. To protect crew members from

massive ejection dangers such as solar flares, an additional enclosure will be created within the

43

main water storage tank with room for two individuals. If a risk of heavy radiation ever arises

during the mission, the crew can safely enclose themselves within this area and have protection

from harmful rays. The water will act as a shield and reduce the magnitude of the radiation crew

members are exposed to.

E. Spacecraft Atmosphere

The internal environment of the spacecraft will simulate the atmospheric makeup of the Earth.

This model will have a composition of 78% nitrogen and 22% oxygen, with the minimal impact

of lesser elements ignored. The oxygen will be produced through the oxygen generation process

detailed in previous sections. Nitrogen will have to be carried on board the spacecraft to meet the

required needs. However, since humans do not use the nitrogen for any biological purposes, the

nitrogen makeup of the overall atmosphere will not fluctuate significantly and will not require

constant replenishing. This means a very limited amount of nitrogen will have to be carried to

simulate this atmosphere. The main reasoning for avoiding a purely oxygen internal environment

is to reduce the danger of flammability. The likelihood of a fire is significantly heightened when

oxygen comprises the majority of the ambient gases; the addition of nitrogen will act to lessen

the overall composition of oxygen and counter this effect. Temperature within the spacecraft will

be controlled by a thermal system dependent on a redundant loop radiator. This will maintain a

suitable living environment for the crew members in an efficient and executable manner.

The condition of the environment will be monitored by an internal piping system that filters the

ambient air. As mentioned earlier, this system will be responsible for separating carbon dioxide

as well as humidity from the air to assist with other life support processes. During this filtering

procedure, the composition of the atmosphere will be recorded and additional nitrogen will be

added or removed if necessary. The ambient temperature will also be monitored within this

system and will control the usage of the loop radiators. Based on the thermal requirements during

certain segments of the mission, either cooling via heat pipes or heating via space heaters will be

present to ensure a suitable living temperature is available for the crew. The requirements vary

based on the spacecraft’s position in the trajectory, as specified earlier in the heat transfer

section, and will dictate the amount heat dumping or heat generation necessary to maintain an

appropriate condition.

44

X. Risk Management

Any failure within the life support system could present a major risk to the mission. As detailed

in previous sections, most aspects of life support rely on subsystems to generate the necessary

output. A mechanical mishap or malfunction within a system could be devastating since the

required daily needs of the crew members might not be fulfilled. For example, both the oxygen

regeneration system and the water recovery system need to function properly for the crew to get

the necessary supplement needed per day. The internal piping system has to perform its duties

otherwise the atmospheric composition of the spacecraft could be compromised. Human error in

regards to tending the Aeroponic system could result in lack of crop production and a decrease in

total food supply. Failure within the misting apparatus could potentially eliminate all crop

production due to a lack of nutrients. Apathy from crew members in reference to daily exercise,

though not a technical risk but still a consideration, could cause medical risks and health issues

to members.

To address these risks, safety margins have been added to most of these life support subsystems.

Crew members will be trained to fix minor issues within subsystems to keep them running full

strength. Extra oxygen, water, hydrogen, and nitrogen will be taken on board to account for

small mishaps and down time during possible repair sessions. Extra food will be brought to

account for both miscues in crop production and fulfilling additional nutritional needs. While all

these precautions have the ability to counter minor failures within the system, any catastrophic

events such as a major system completely ceasing to operate could result in loss of life. While

these subsystems have gone through extensive testing for survivability, the possibility of failure

still exists and is something the crew members need to accept when embarking on this mission.

The power system has very few risks, the only notable ones being the deployment mechanism as

well as the Lithium Polymer batteries. If the deployment mechanism were to fail, the spacecraft

would essentially be without power. It is imperative that the mechanism works flawlessly and

has a level of redundancy. The Lithium Polymer batteries also pose a risk because they are

relatively new to space travel. There is always the possibility of LiPo batteries overheating,

catching fire, exploding, etc. It would not only be an issue that the spacecraft is subject to

undesirable conditions but it also would be without power. Part of this issue is addressed by the

inclusion of a battery contingency so if something were to go wrong with a small portion of

batteries, the issue could resolved. Additionally, the batteries are dispersed circularly throughout

the spacecraft which has numerous benefits. One benefit will be the dispersion of the weight of

the batteries rather than a larger concentrated mass. Furthermore, if one of the batteries happened

to explode or catch fire the issue would be concentrated to one area and would not affect all of

the batteries on the spacecraft.

45

XI. References

[1] Rankin, Daniel A. "A Model of the Cost of Software Development for the Apollo Spacecraft

Computer." http://www.ibiblio.org/apollo/hrst/archive/1728.pdf

[2] Wertz, James R., Jeffery John. Puschell, and David F. Everett. "11. Cost Estimating." Space

Mission Engineering: The New SMAD. N.p.: Microcosm, 2011. 289-318. Print.

[3] Jones, Harry. "Equivalent Mass versus Life Cycle Cost for Life Support Technology

Selection." Society of Automotive Engineers (2003): n. pag. Web.

[4] Moonish R. Patel, James M. Longuski, and Jon A. Sims, "Mars Free Return Trajectories,"

Journal of Spacecraft and Rockets, vol. 35, no. 3, pp. 350-354, May–June 1998.

[5] Corum, Battiste, Liu, and Ruggles. “Basic Properties of Reference Crossply Fiber

Composite.” Oak Ridge National Laboratory-Lockheed Martin.

http://web.ornl.gov/~webworks/cpr/v823/rpt/106099.pdf

[6] Milos, F. S., and Y. K. Chen. "Ablation and Thermal Response Property Model Validation

for Phenolic Impregnated Carbon Ablator." Journal of Spacecraft and Rockets 47.5 (2010): n.

pag. Web.

[7] Keesee, John. “Spacecraft Thermal Control Systems.” Massachusetts Institute of Technology.

Ppt. <http://ocw.mit.edu/courses/aeronautics-and-astronautics/16-851-satellite-engineering-fall-

2003/lecture-notes/l23thermalcontro.pdf>.

[8] United States. NASA. Apollo Command Module Overview. Web.

<http://www.hq.nasa.gov/alsj/CSM06_Command_Module_Overview_pp39-52.pdf>.

[9] Hartsfield, Carl. “More Space Propulsion Information-Sizing and Materials.” Class lecture

material. Ppt.

[10] “Xenon Element Facts.” Chemicool. <http://www.chemicool.com/elements/xenon.html>.

[11] Krikorian, Y.Y.; Emmons, D.L.; McVey, J.P., "Communication coverage and cost of the

deep space network for a Mars manned flyby mission," Aerospace Conference, 2005 IEEE , vol.,

no., pp.1670,1677, 5-12 March 2005 doi: 10.1109/AERO.2005.1559460

[12] http://deepspace.jpl.nasa.gov/advmiss/index.html

[13] “Hybrid Inflatable Antenna (HIA).” ILC DOVER. Web. 14 Nov 2013. Retrieved from

http://www.ilcdover.com/Hybrid-Inflatable-Antenna-HIA/

[14] Brammer, Paul. “uTx-300 Ka-Band Transmitter Unit.” Email to Nick Filipkowski. 12 Nov.

2013

46

[15] "Heritage Receiver/Downconverter Line Up." . NEC TOSHIBA Space Systems. Ltd.. Web.

14 Nov 2013.

<http://www.nec.com/en/global/solutions/space/satellite_communications/images/Ka-Ku-

band_RCVR.pdf>.

[16] "μTx-300 Ka-Band Transmitter." . Space Micro, Inc.. Web. 14 Nov 2013.

<http://www.spacemicro.com/pdfs/KA-Band v5.0.pdf>.

[17] Peterson, L. (2009). Environmental Control and Life Support System (ECLSS) [PowerPoint

slides]. Retrieved from

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20090029327_2009028591.pdf

[18] Carrasquillo, R. (2013). ISS Environmental Control and Life Support System [PowerPoint

slides]. Retrieved from http://astronautical.org/sites/default/files/issrdc/2013/issrdc_2013-07-17-

1600_carrasquillo.pdf

[19] "NASA Targets Water Recycling System for Rapid Development." NASA. NASA, n.d.

Web. 21 Nov. 2014.

<http://www.nasa.gov/centers/ames/news/2013/WaterRecyclingSystem_7_Feb_2013.html#.Uxz

bVPldWrk>.

[20] Carter, Layne. "Status of the Regenerative ECLS Water Recovery System." American

Institute of Aeronautics and Astronautics (n.d.): n. pag. Print.

[21] "Inflatable Aeroponic System." NASA. NASA, 16 Nov. 2009. Web.

<http://www.nasa.gov/offices/ipp/centers/kennedy/success_stories/Inflatable_Aeroponic_System

_BBlinds.html>.

[22] I, Nir. "Growing Plants in Aeroponics Growth System." (n.d.): n. pag. Web

[23] Cucinotta, Francis A. "SPACE RADIATION CANCER RISK PROJECTIONS FOR

EXPLORATION MISSIONS: UNCERTAINTY REDUCTION AND MITIGATION." (2001):

n. pag. Web

[24] Rojdev, Kristina, and Eric Christiansen. "Advanced Multifunctional MMOD Shield:

Radiation Shielding Assessment." AIAA (n.d.): n. pag. Web.

[25] "Understanding Space Radiation." National Aeronautics and Space Administration (2002):

n. pag. Web.