a stronautical reconnaissance expedition spacecraft to...
TRANSCRIPT
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
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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
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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
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[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-
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[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
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[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-
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[19] "NASA Targets Water Recycling System for Rapid Development." NASA. NASA, n.d.
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<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.
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_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:
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[25] "Understanding Space Radiation." National Aeronautics and Space Administration (2002):
n. pag. Web.