Romulus Project Mission Report from the Moon to Mars

ROMULUS PROJECT MISSION REPORT:
FROM THE MOON TO MARS
SYSTEM DEFINITION REVIEW
LETOURNEAU UNIVERSITY
2017

COVERSHEET
ROMULUS PROJECT
MISSION REPORT
LETOURNEAU UNIVERSITY
MEGR 4993
MEGR 5993
MEGR 6993
SUBMITTED: 25 APRIL 2017

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PROJECT INTRODUCTION
Since the dawn of the Roman Empire, humans have intensely pondered the celestial bodies.
When the Romans looked up at the night sky, they saw their greatest heroes. Instead of a
constellation of stars, they saw the hunter god Orion. Instead of a planet, they saw the god of
beauty – Venus. Instead of a red dot, they saw the planet Mars, which was named after the
Roman war god of the same name.
In ancient days, men looked to the heavens and saw their mythological heroes. Soon, men will
look towards the heavens and see our actual heroes – men and women from planet Earth on their
way to visit Mars. The first manned mission is scheduled to occur in the 2030s. This proposal,
Project Romulus, is a conceptual plan for the spacecraft that will carry these heroes on their
journey to the Red Planet.
The Roman god Romulus was the founder of Rome. The Project Romulus team hopes that this
mission will be the beginning of a colony on Mars – one that will endure like the Roman empire.
Additionally, Romulus was the son of the war god Mars, making the name doubly suitable for a
manned mission to the planet Mars.
The initial conceptual plan is to use Earth’s Moon as a staging point for the Romulus mission. As
many of the mission’s supplies as possible will be manufactured or mined on the Moon –
primarily structural components and fuel. Once manufacturing is complete and all required
materials are obtained, the spacecraft will be launched from the Moon. This concept and the
spacecraft design has been further refined and developed, culminating in this report.

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REPORT INTRODUCTION
This work presents the design that has been developed for safely transporting nine crew members
from the lunar surface to Martian orbital insertion. Included are reports from each of the major
systems engineering teams, and a project overview and report from the Project Management
team. Together, this document describes the spacecraft’s design and specifications to date.
Systems have been selected for each major project requirement, and thermal, electrical, mass,
and volume footprints have been estimated for each system. These characteristics have then been
used to design the electrical and thermal systems necessary for proper function of the craft and
protection of the crew. Additionally, supplies such as food, water and oxygen, have been
considered, and a multistage propulsion system has been designed for transiting the crew module
to the Martian orbit.
This report fulfills the requirements of a system definition review (SDR) as defined by NASA’s
procedural requirements. An SDR is accomplished prior to the detailed design. The SDR ensures
that the all technical requirements and designs are at an adequate level of maturity such that
development of new technologies can begin. In order to accomplish this, the SDR establishes
various success criterion. These include system technical requirements with sub-system flow
down and resource allocation within mission parameters, a sound requirements process, credible
technical approach, updated technical plans, and completed tradeoff studies, which show that the
technical plans have been updated to reflect the higher level of maturity when compared to
previous technical reviews. In addition to these, the SDR also assesses mission(s) development,
adequate planning for new technology, and mission operation concept which is consistent with
proposed mission concepts and aligned with the mission requirements. Lastly, this review

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evaluates risks that have been assessed through design studies and presents a plan to handle those
identified risks and to ensure health and safety of the crew throughout the mission.

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TABLE OF CONTENTS
...............................................................................................................................................Page
Coversheet........................................................................................................................................ i
Project Introduction ........................................................................................................................ii
Report Introduction........................................................................................................................iii
Table of Contents............................................................................................................................ v
Project Management Team ............................................................................................................. 1
Overview............................................................................................................................. 2
System Integration .............................................................................................................. 3
System Effects and Coupling................................................................................... 3
Design Sequence ..................................................................................................... 5
Space Craft Geometry......................................................................................................... 7
Project Timeline................................................................................................................ 13
Concluding Statements ..................................................................................................... 16
References......................................................................................................................... 17
Appendix A: Iterative Design Process Flow Chart........................................................... 18
Appendix B: Itemized Mass Estimates............................................................................. 19
Appendix C: Itemized Dimensions Estimates .................................................................. 20
Appendix D: Itemized Electrical Power Loads Estimates................................................ 21
Appendix E: Itemized Heat Loads Estimates ................................................................... 22

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Propulsion Team........................................................................................................................... 23
Overview........................................................................................................................... 24
Trade Studies .................................................................................................................... 24
System Design Studies...................................................................................................... 25
System Configuration ....................................................................................................... 26
Fuel Tank Configuration....................................................................................... 26
References......................................................................................................................... 28
Appendix A: Delta-v Calculations.................................................................................... 30
Direct Lunar Transfer........................................................................................... 30
Earth Oberth Effect Transfer................................................................................ 31
For Both Direct and Earth Oberth Effect Transfer .............................................. 33
Delta-v Totals........................................................................................................ 34
Delta-v Difference................................................................................................. 34
Communications and Data Team.................................................................................................. 35
Overview........................................................................................................................... 36
Trade Studies .................................................................................................................... 36
Communications:.................................................................................................. 36
Data....................................................................................................................... 38
Communications ................................................................................................... 39
Data....................................................................................................................... 41

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Communications ................................................................................................... 43
Tracking Coverage: .............................................................................................. 45
Data....................................................................................................................... 46
References......................................................................................................................... 49
Appendix A: Radiation Characteristics of CI-1HN89 NAND Flash Memory (from
datasheet) .......................................................................................................................... 50
Electrical Power Team.................................................................................................................. 51
Overview........................................................................................................................... 52
Trade Studies .................................................................................................................... 52
Solar Panels.......................................................................................................... 56
Li-Ion Batteries..................................................................................................... 57
Power Control Distribution Unit (PCDU) ........................................................... 58
Power Cables........................................................................................................ 59
References......................................................................................................................... 62
Appendix A: Trade-Off Studies........................................................................................ 64
Appendix A.1: Power Source Trade-Off Study Results......................................... 64
Appendix A.2: Power Storage Trade-Off Study Results ....................................... 65
Appendix B: Solar Energy Flux........................................................................................ 66
Appendix C: Power vs Solar Panel Area .......................................................................... 68

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Appendix D: Power vs Solar Panel Mass ......................................................................... 69
Appendix E: Batteries....................................................................................................... 70
Thermal Control Team.................................................................................................................. 72
Overview........................................................................................................................... 73
Trade Studies .................................................................................................................... 73
Insulation .............................................................................................................. 73
Heat Generation.................................................................................................... 74
Electrical System Cooling..................................................................................... 75
Heat Transfer........................................................................................................ 75
Heat Transfer Coolant.......................................................................................... 76
Heat Rejection....................................................................................................... 77
Cabin Thermal Control......................................................................................... 78
Insulation .............................................................................................................. 79
Heat Generation.................................................................................................... 81
Heat Transfer........................................................................................................ 86
Heat Rejection....................................................................................................... 87
Insulation .............................................................................................................. 90
Heat Generation.................................................................................................... 90

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Heat Transfer........................................................................................................ 91
Heat Rejection....................................................................................................... 92
References......................................................................................................................... 95
Environmental Control and Life Support Systems ....................................................................... 97
Overview........................................................................................................................... 98
Trade Studies .................................................................................................................... 98
Waste Processing.................................................................................................. 98
Atmospheric Revitalization................................................................................. 100
Oxygen Generation ............................................................................................. 102
Fire Detection and Suppression.......................................................................... 104
Consumables Storage.......................................................................................... 108
System Design Studies.................................................................................................... 108
Waste Processing................................................................................................ 108
Fire Suppression................................................................................................. 111
Human Heat Load............................................................................................... 113
System Configuration ..................................................................................................... 113
Waste Processing................................................................................................ 113

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Fire Detection and Suppression.......................................................................... 117
Concluding Statements ................................................................................................... 118
References....................................................................................................................... 120
Appendix A: Trade Studies............................................................................................. 123
Appendix A-1: Waste Processing Trade Study ................................................... 123
Appendix A-2: Atmospheric Revitalization Trade Study .................................... 124
Appendix A-3: Oxygen Generation Trade Study ................................................ 125
Appendix A-4: Fire Detection and Suppression Trade Study............................. 126
Appendix A-5: Consumables Storage Trade Study............................................. 127
Structures Team .......................................................................................................................... 128
Overview......................................................................................................................... 129
Trade Studies .................................................................................................................. 131
System Design Studies.................................................................................................... 132
Radiation Shielding............................................................................................. 132
System Configuration ..................................................................................................... 134
Double Bubble .................................................................................................... 134
Pressure vessel calculations ............................................................................... 135
Stage attachment................................................................................................. 136
Stage 1-2 connection:.............................................................................. 136
Stage 2-3 connection:.............................................................................. 136
Stage 3-Capsule connection:................................................................... 137

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Concluding Statements ................................................................................................... 137
References....................................................................................................................... 138
Appendix A: Initial Concepts ......................................................................................... 139
Appendix B: Double Bubble Design .............................................................................. 141
Appendix C: Mass Estimates and Zvezda Comparison.................................................. 144

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ROMULUS MISSION REPORT
PROJECT MANAGEMENT TEAM
AARON J. CONRAD
JOSH HOOKS
JUDAH RUTLEDGE
CLIFF WHITE
25 APRIL 2017

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TEAM: PROJECT MANAGEMENT
CONRAD, HOOKS, RUTLEDGE, WHITE
DATE: APRIL 25, 2017
OVERVIEW
The Project Management report discusses the parallel design strategies used, describes
system integration challenges, and presents an updated and comprehensive project timeline. The
parallel system design strategy ensures completion of technical requirements, system to sub-
system requirement flowdown, and allocation of system resources. The operational concept is
also demonstrated to be consistent with the mission goals. Additionally system integration
demonstrates how the system loads were determined while considering the complex interplay
between various subsystems.
In addition to managing system loads, this report presents the final geometry based on the
allocation of mass and the physical systems external to the spacecraft, such as solar arrays and
radiator arrays and presents the first estimates of the dimensions and mass of the final craft.
Finally, the project management team has created a comprehensive project development
timeline. This timeline allows for the time needed to develop the many new technologies
required for the Romulus project, while still meeting the launch window requirements. The
timeline also demonstrates that the operational concept is consistent with the mission goals.

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SYSTEM INTEGRATION
System integration ensures that the sub-systems flow together and verifies that available
resources are adequate for system and personal needs. The space craft systems were developed in
parallel to reduce the overall design time. To organize the design order, the various systems were
evaluated to determine which system parameters were strongly, and which were loosely coupled.
This allowed the integration team to determine the most effective order of design.
System Effects and Coupling
In order to perform this analysis, mass, electrical consumption, and thermal generation
were chosen as the primary spacecraft loads, and each system’s effect on these loads was
considered. A system affects a load when the system needs change the load. A system is coupled
with a load when the system not only affects the load, but the load also affects the system. First,
loosely coupled systems and systems with a low effects identified (Figure 1). Based on input
from team leads, it was estimated that propulsion and structures had little input on the thermal
load, since the thermal loads generated by the propulsion system are of short duration, and each
stage is ejected after use. Similarly, the overall electrical load of propulsion and structures is
small in comparison to the requirements of the other teams (ECLSS and Thermal, being the
foremost). Finally, a loose coupling exists between the craft’s mass and structural design.
Increasing the mass could impact structural calculations, and require a redesign. However, this
redesign is unlikely to, in turn, change the mass significantly.

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Figure 1: Low dependence and loosely coupled systems.
After the loosely coupled systems were identified, the systems that strongly affected or
were tightly coupled with loads were considered. It was self-evident that thermal and electrical
were tightly coupled with the heat and power needs. Additionally, Propulsion and mass are
tightly coupled, since fuel comprises most of the mass of the finalized craft, and the amount of
fuel necessary depends on the mass of the finalized crew module. Additionally, it can be seen
that all systems affect the mass, that ECLSS, Comm./Nav, and Electrical primarily affect Heat
loads, and that ECLSS, Comm./Nav. and thermal are the primary inputs for Electrical loads.
Figure 2: Strong effects and tightly coupled systems.
Propulsion Structures
Thermal
Electrical
Mass
Com/Nav
ECLSS
Heat
Electricity
Propulsion Structures
Thermal
Electrical
Com/Nav
ECLSS
Mass
Heat
Electricity

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Design Sequence
In order to select the order of the system design. Systems that were not coupled with any
space craft characteristics were finalized first. These were Comm./Nav, ECLSS, and Structures.
The load characteristics of these systems, once determined, were fed to the Thermal and
Electrical teams, which designed their systems to handle the required loads. After these teams
finalized their designs, the total mass of the crew module was able to be determined, and the
propulsion needs and total craft mass were calculated. In order to mitigate time losses because of
waiting on other teams, various tasks were assigned to each team in parallel while the system
loads were being determined, such as calculating the Delta V, for propulsion, or deriving a
relationship between system mass and capacity for the Thermal and Electrical teams. The
complete breakdown of these tasks is shown in Figure 3, which is a timeline of the schedule for
each major milestone during the system definition phase. Each team was responsible for the
system items in their respective columns, which could be completed only after consideration of
the pre-requisites feeding into the task. The flow-chart nature of the figure demonstrates the
complex interdependencies of the systems and their developmental timelines.
As an example of the interdependent nature of the design process, consider the electrical
system column. The electrical team’s first decisions during the system definition phase were to
determine the solar energy available as a function of distance from the sun and panel
degradation. Additionally, the team modeled the electrical systems total mass as a function of
power generated. These parameters then influenced the electrical system’s thermal loads, the

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Figure 3: System Definition Design Schedule

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preliminary system configuration for the electrical system, and the final system design. Each step
(except the first) also depended on the results of other system team’s work. At this point in time,
the design phase was cut short. The repercussions of the system loads on the loosely coupled
systems with lower overall effects were not considered. Completing the design process requires
further modeling of these loosely coupled systems, which is more iterative in nature. Appendix A
contains flow chart highlighting the iterative nature of the design process.
SPACE CRAFT GEOMETRY
The finalized space craft design is shown in Figure 4. The living quarters of the craft are
comprised of a Command Module, and a Crew Module. The crew module is larger and houses
the sleeping and living quarters of the Romulus crew, while the command module contains all of
the space craft controls, communication systems, and data management. The command module
can be isolated from the crew module in event of an emergency, providing the crew a safe haven
with full access to craft controls and FDIR systems.
Both the crew module and the command module have observation windows facing the
front and aft of the craft. The crew module and support systems comprise 85,000 kilograms. The
total frontal area of the craft in flight is 52 x 69 meters including the deployed solar wings and
radiation panels. (Figure 5)
When integrating the craft, it was decided to place the last stage upside down with respect
to the surface of the moon (Figure 6). This places the last stage and the craft modules in the
proper orientation for Mars orbital insertion from the outset of the mission. Since the maximum
G loading experienced by the crew members during orbital insertion to Mars is 0.44 Earth gs,

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crew health and performance would not be adversely affected by experiencing this G-loading in
an inverted position.
The craft has 3 stages and a diameter of 16 meters and a height of 120 meters on the
launch pad. The total craft mass with propellant for the 3 stages is 1.6 million kilograms. The
first stage is used to lift off from the moon and enter Earth’s orbit, the 2nd
stage ejects from
Earth’s orbit and places the craft on an interorbital trajectory to Mars, and the 3rd
stage performs
a Mars orbital insertion maneuver, marking the end of mission.

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Figure 4: Finalized geometry for the Romulus craft Modules
entering orbit around Mars.

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Figure 5: Romulus Craft Footprint from the front with deployed radiation
panels and solar arrays.
68.5meters
52.4 meters

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Figure 6: Romulus Craft in launch pad configuration (launch pad not shown).
68.5meters
16 meters

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PROJECT TIMELINE
The timeline that follows is based on the NASA project Life Cycle as shown in Figure
5.2 of NASA Procedure and Guidelines NPR 7123.1A – Chapter 5. This entire report fulfills the
requirements of the System Definition Review as outlined in chapter 5 of NPR 7123.1A. After
the System Definition Review, specific dates for technological development will be set inside the
box outlined. Also during this time, there will be dates for further concept development and
testing.
The preliminary design review as defined by NASA is scheduled to be completed on June
27, 2022. This design review includes but is not limited to: fully assessed risks, adequate
technical margins, developed new technologies, and a technically sound operational concept.
After this date, the final design phase will begin. During this phase, the design will be finalized
and the fabrication of new equipment will begin.
The critical design review will be conducted midway through this phase. This review will
include: a detailed design, interface control documents, confidence in the baseline product,
product verification, comprehensive testing approach, adequate technical and programmatic
margins, and understanding of risks to mission success. At the end of the phase, a System
Integration Review will be conducted which is scheduled for March 27, 2025. This review looks
at how the system interact with each and ensures that the interactions between these systems are
technically and operationally sound.
The next step will begin the craft assembly, integration, and testing. During this phase, an
operational readiness review will be conducted. The operational readiness review ensures that
the equipment is ready to be put into operational status and that any anomalies and waivers have
been closed. Lastly, this phase will end with final craft assembly. Final testing will be completed

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and the flight readiness review will be completed one week prior to the scheduled launch date of
January 3, 2029. A flight readiness review makes sure that the flight vehicle is read for flight, all
interfaces are checked, any open items are determined to be acceptable, hardware is deemed safe,
all safety items have been addressed, and flight and recovery environmental factors are within
constraints. This review ensures the system is ready for launch.

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CONCLUDING STATEMENTS
The Project Management team has effectively organized and managed design
deliverables for the completion of the system definition phase of the Romulus Project. Parallel
design was accomplished by analyzing the craft for tightly and loosely coupled systems and
loads. This analysis served to determine major design markers and determine the order of design.
System modeling of the final system mass as a function of the system loads was also used to
accelerate the design process.
The first estimates of the craft’s mass and dimensions were determined, as well as the
orbit transfer being pursued and the individual engine stages. From this system definition, the
project is ready to move forward into refining and developing the technology needed to complete
the design and assembly of the craft. Major project milestones include the Preliminary Design
Review scheduled for 2022, The Critical Design Review in 2024, and the beginning of the
Romulus craft construction, in 2025. After the craft construction and integration testing, the
Flight Readiness Review and Launch are slated for January 2029.

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REFERENCES
Boeing. "Module J: Trade Studies". 2017. Presentation.
“Mars Science Laboratory Mission Profile.” Spaceflight 101. N.p., n.d., Web. 09 March 2017
"NASA Procedures and Guidelines." Section NPR 7123.1A, Chapter 5, NASA. NASA, 26 Mar.
2007. Web. 23 Apr. 2017.
Pisacance, Vincent L. Fundamentals of Space Systems: Second Edition. Oxford, 2005.
Uhlig, Thomas, Florian Sellmaier, and Michael Schmidhuber.. Spacecraft Operations. 1st ed.
Vienna, New York, Dordrecht, London: Springer, 2015. Verlag Gmbh, 2016. Print.

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APPENDIX A: ITERATIVE DESIGN PROCESS FLOW CHART

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APPENDIX B: ITEMIZED MASS ESTIMATES
Grand Total:
System: Item: Quantity: Mass (kg): Net Mass: 926976.47
Comms Array 1 15.00 15.00
Data System 1 1.77 1.77 Total:
16.77
Li-Ion Battery 7 175.00 1225.00
Power Cable 160934 m 0.012 kg/m 1950.00
Power Control Distribution Units (PCDUs) 4 30.00 120.00 Total:
Solar Panel Wings 2 700.00 1400.00 4695.00
Atmosphere (20% oxygen, 80% nitrogen) 280 1.20 335.50
Backup breathing masks 9 2.25 20.25
Carbon dioxide delivery tubes & nozzles 1 3.00 3.00
Carbon dioxide storage (tank + gas) 2 4.00 8.00
Electrolytic Converter 1 75.00 75.00
Food (days worth) 213 1.77 377.01
HEPA filters 10 2.00 20.00
High Pressure Storage Tanks 4 105.00 420.00
Humidity separator fans 10 0.50 5.00
Medicine 213 0.01 2.13
Metal-Oxide scrubbers 10 1.10 11.00
Photoelectric detectors 15 0.80 12.00
Solid Fuel Oxygen Generators 18 1.50 27.00
Waste collection system 2 300.00 600.00
Water (days worth) 300 2.42 726.00
Water reclamation system 1 170.00 170.00 Total:
Water-mist dispersion extinguishers 5 3.50 17.50 2829.39
1st Stage Tank 1 58388.00 58388.00
2nd Stage Tank 1 7080.00 7080.00
LH2 1 637320.00 637320.00
LOX 1 115429.00 115429.00
Rocketdyne J-2 4 1578.00 6312.00 Total:
Zero Boil-off Tank 1 2926.00 2926.00 827455.00
Big Bubble 1 43241.52 43241.52 Total:
Little Bubble 1 27240.45 27240.45 70481.98
Ammonia (L) 1170 0.64 744.40
Cabin Heat/Cool 1 82.00 82.00
Electric Foil Heater 30 0.04 1.23
Heat Exchangers 12 40.00 480.00
Loops 4 0.2 kg/m 200.00
MLI 1 155.00 155.00
Pumps 12 350.00 4200.00
Radiator Panels 9 1200.00 10800.00
Radiator Rotors 3 440.00 1320.00
Tank Assembly 4 700.00 2800.00 Total:
Triol (L) 700 1.02 715.70 21498.33
Communications,
Navigation, and
Data
Thermal Control
Structure
Propulsion
Environmental
Control and Life
Support Systems
Electrical Power

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APPENDIX C: ITEMIZED DIMENSIONS ESTIMATES
System:Item:Quantity:Dimensions(m):NetVolume(m^3):Notes:
CommsArray115Outside,antennasmustbefacingEarth
DataSystem10.000185Insidethecraft,climatecontrolled
Li-IonBattery73.2x0.34x0.42.9Inside
PowerCableNegligible160934,0.0015D0.28Inside
PowerControlDistributionUnits(PCDUs)40.6x0.345x0.1950.162Inside
SolarPanelWings227x5x0.0513.5Outside
BackupBreathingMasks90.5x0.37D0.50Inside(Stationedinemergencyfallbackarea)
ElectrolyticConverter12.55(2x1.5x0.85)2.55Inside(withotherrackedsystems)
FireExtinguishers50.5x0.5D0.50Inside(Distributedinhabitableareas)
HEPAFilters101x.1x.10.01Inside(spacedthroughoutcraftindifferentlocations)
HighPressureStorageTanks41.09(1x.35D)4.36Inside(withotherstoragetanks)
SFOG180.3x0.08D0.0271
WasteCollectionSystem21x2x.51.00Inside(inseperatelocations)
WasteStorageTanks2.3x.65D0.10Inside(withotherstoragetanks)
WaterRecoverySystem1(2x1.5x0.85)2.55Inside(withotherrackedsystems)
WaterStorage11.3x1D1.02Inside(withotherstoragetanks)
LH21stStage135mx2.66m^293.1
LH22ndStage118mx1.57m^228.26
LH23rdStage119mx1.1m^220.9
LOX1stStage12.66
LOX2ndStage11.57
LOX3rdStage11.1
BigBubble14.571D400
LittleBubble13.63D200
Ammonia(L)11701.17InsidePiping
CabinHeat/Cool10.565x1.134x0.1460.09354366InsideCrewCabin
ElectricFoilHeater30NegligibleNegligibleInside
HeatExchangers120.64x0.53x0.20.81408Outside
Loops40.0095x0.0095x10000.361Inside
MLI120x20x0.028Outside
Pumps121.8x1.3x0.9125.5528Inside
RadiatorPanels93.1x0.1x2364.17Outside
RadiatorRotors31.7x1.4x1.39.282Outside
TankAssembly43.6x2.1x2.266.528Inside
Triol(L)7000.7InsidePiping
ThermalControl
Communications,
Navigation,and
Data
ElectricalPower
Environmental
ControlandLife
SupportSystems
Propulsion
Structure

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APPENDIX D: ITEMIZED ELECTRICAL POWER LOADS ESTIMATES
System:Item:Quantity:Input(W):NetPower:Notes:
Amplifiers2100200
DataSystem14848
High-gainAntenna1100100Onlyoneantennaisusedatatime,so
Low-GainAntenna2100200truemaxhereisonly100Watts
RAD5545SpaceVPX
single-boardcomputer635210
Normaloutputisexpectedtobeabout
halfofthemaximum
Transponder200TransponderusespowerfromAntenna
Li-IonBattery7StoresStores|PowergeneratednearEarth.
PowerCableNegligibleTransfersTransfers|Travellingawayfromsun,safety
PowerControlDistributionUnits(PCDUs)4TransfersTransfers|factor,andradiationdemands
SolarPanelWings2-24500-49000foralargerpowergeneration
ElectrolyticOxygenGenerator112001200
HEPAFilters101001000
HumiditySeparator1030300
PhotoelectricFireDetectors151.522.5
WasteCollectionSystem2350700
WaterReclamationSystem1700700
PropulsionZeroBoil-offTank1100100
Ammonia(L)117000
CabinHeat/Cool1600600
ElectricFoilHeater301203600
HeatExchangers1200
Loops400
Pumps126007200
RadiatorPanels900
RadiatorRotors31030
TankAssembly400
Triol(L)70000
ThermalControl
Environmental
ControlandLife
SupportSystems
ElectricalPower
Communications,
Navigation,and
Data

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APPENDIX E: ITEMIZED HEAT LOADS ESTIMATES
System:Item:Quantity:Output(W):NetHeat:Notes:
CommsArray13030Normaloutputis
DataSystem14848expectedtobeabout
RAD5000SeriesComputer635210halfofthemaximum
Li-IonBattery7100700
PowerCable160934m50005000
PowerControlDistributionUnits(PCDUs)44001600
SolarPanelWings2500010000
CrewHeatLoad91.1610.44
ElectrolyticOxygenGenerator1200200
HEPAFilters1010100
HumiditySeparator10330
PhotoelectricFireDetectors15115
WasteCollectionSystem288176
WaterReclamationSystem1175175
Ammonia(L)117000
CabinHeat/Cool16060
ElectricFoilHeater3000
HeatExchangers1200
Loops400
Pumps1260720
RadiatorPanels900
RadiatorRotors326
TankAssembly400
Triol(L)70000
Communications,
Navigation,and
Data
ElectricalPower
Environmental
ControlandLife
SupportSystems
ThermalControl

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PROPULSION TEAM
TEAM LEAD: DAVID RING
SARAH COPELAND
ELLI KEENER
BEN KEM
CAT NIX
25 APRIL 2017

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PROJECT ROMULUS REPORT
TEAM: PROPULSION TEAM
COPELAND, KEENER, KEM, NIX, RING
DATE: 25 APRIL 2017
OVERVIEW
The Propulsion Team was responsible for three key areas: to ensure efficient and successful
launch from the lunar surface, transit to Mars, and insertion into a Martian orbit. To do so, the
team compared available options for engines, fuel types, fuel storage, and trajectory.
TRADE STUDIES
Engines were compared with four key criteria: thrust, specific impulse, safety, and TRL. These
criteria were chosen because an engine needs to be efficient, powerful, manufacturable, and safe.
Thrust and specific impulse were weighted the highest because it is crucial that the engine is able
to launch the spacecraft and to do it efficiently. Safety was chosen because the spacecraft will
have humans on board, and human life is valuable. Our projected launch date is January 3rd,
2029, so the technology that we choose must be ready in time for launch. TRL was chosen as the
fourth criteria for this reason. Based on these comparisons, the RL-10 and Rocketdyne J-2 are
the best candidates for engine selection for the Romulus Mission. For the full weighted
evaluation, see the table below:

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There are three stages with different engines used for each stage. The first stage, leaving
the moon’s surface, will use three Rocketdyne J-2 engines. These three engines give a thrust to
weight ratio of 1.23 when lifting off from the moon. The second stage will use one Rocketdyne
J-2 engine to leave the moon and eject from orbit around the earth. The third stage will use four
RL-10’s to insert the spacecraft into a Martian orbit. The engines for the second and third stages
were chosen in order to keep burn times under ten minutes.
There were two options when choosing the trajectory, a direct ejection from low lunar
orbit and ejecting into an elliptical orbit around Earth and using the advantage of the Oberth
effect. Between the two, the Oberth effect path offered a 345 m/s savings on the delta-v required,
but added 6 days to the transit time. This trajectory was chosen because the mass savings in fuel
were more significant than the added the mass of the extra supplies. For more details on the
delta-v calculations, see Appendix A. Additionally, the team chose to eject using into a
Hohmann transfer due to its time and fuel efficiency.
SYSTEM DESIGN STUDIES
When calculating interplanetary trajectories without computing software, a few
assumptions had to be made. The propulsion team assumed that Earth and Mars orbited in

26
circular and coplanar orbits around the sun. The team felt these were reasonable assumptions
seeing as Earth has an eccentricity of 0.017 and Mars has an eccentricity of 0.093. Mars also has
an inclination of 1.85 degrees relative to Earth. These were used for the trajectory calculation
(see Appendix A) that lead to the decision to eject into an elliptical orbit around Earth and then
to eject to Mars.
SYSTEM CONFIGURATION
Fuel Tank Configuration
The fuel tanks configuration will include a first stage of a cluster of engines to escape
from the moon's gravity. This configuration will produce dimensions of approximately 16m in
diameter by 34.8m long. Second a single engine will be used to escape the orbit of the
moon/earth, producing dimensions of 10m in diameter and 18.3m long. The final stage for mars
insertion will include a zero boil off tank in order to keep the needed fuel in cryo. This tank is
able to keep cryogenic fuels at the temperatures necessary for no boil off to occur, eliminating
the need for extra fuel to be brought as compensation. Dimensionally the final stage will be 6m
diameter and 19m long. In order to enable easier structural mounting, circular tanks will be
used. Due to short length of time between lift off and ejecting from the moon’s orbit no zero boil
off tank will be needed for the first or second stages. Enough fuel will be added to compensate
for the boil off between takeoff and escape from the moon’s orbit. Burns for the second and third
stages will be optimized to keep burn times under ten minutes. Finally the three engines for lift
off will produce a combined thrust to weight ratio of 1.23 on the moon.
In conclusion, the systems designed will allow the spacecraft to lift off from the lunar
surface, eject to a transfer orbit, and insert into low Martian orbit. The systems will safely

27
transport the crew to Mars in a reasonable amount of time while providing a safe transport. The
propulsion system uses fuel obtained from lunar resources which removes the need for fuel to be
brought from Earth. It was designed to keep the burn times under ten minutes ensuring the burns
performed will be efficient and put the craft under minimum vibrational stress and mechanical
failures are unlikely. The total transit time was 270 days, which is reasonable given the fuel
savings by taking the more efficient trajectory.

28
REFERENCES
Plachta, D., & Kittel, P. (2003, June). An Updated Zero Boil-Off Cryogenic Propellant Storage
Analysis Applied to Upper Stages or Depots in an LEO Environment. Retrieved from
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20030067928.pdf
Aerojet Nuclear Systems Company. Performance/Design And Qualification Requirements For
Engine, NERVA, 75K, Full Flow. 1970. Print. NERVA.
Aerojet Rocketdyne. Aerojet Rocketdyne RL10 Propulsion System. 2016. Print.
Aircraft And Rocket Propulsion. 1st ed. Stanford University, 2017. Web. 6 Mar. 2017.
Erichsen, Peter. Spacecraft Propulsion: A Brief Introduction. 2005. Print.
Friesen, Larry Jay. "Moon Miners' Manifesto: Lunar Aluminum And Oxygen Propellants".
Asi.org. N.p., 1996. Web. 6 Mar. 2017.
"Hybrid Rocket Engines - Copenhagen Suborbitals". Copenhagen Suborbitals. N.p., 2017. Web.
6 Mar. 2017.
J., D. R., Damaren, C., & Forbes, J. R. (2013). Spacecraft dynamics and control: an introduction.
Chichester: John Wiley.
Mihaila-Andres, Mihai and Paul Virgil Rosu. Thermo-Gas Dynamic Analysis Of Upper-Stage
Rocket Engine Nozzle. International Conference of Scientific Paper, 2015. Print.
NASA. An Historical Perspective Of The NERVA Nuclear Rocket Engine Technology Program.
NASA, 1991. Print.
NASA. Large-Scale Demonstration Of Liquid Hydrogen Storage With Zero Boiloff For In-Space
Applications. NASA, 2010. Print.
NASA. Mastering Cryogenic Propellants. NASA, 2013. Print.

29
Plachta, D.W., W.L. Johnson, and J.R. Feller. Cryogenic Boil-Off Reduction System Testing.
NASA. Print.
Zakirov, Vadim et al. "Nitrous Oxide As A Rocket Propellant". Acta Astronautica 48.5-12
(2001): 353-362. Web. 5 Mar. 2017.
Sovey, J. S., Rawlin, V. K., and Patterson, M. J.: "Ion Propulsion Development Projects in U. S.:
Space Electric Rocket Test 1 to Deep Space 1." Journal of Propulsion and Power, Vol.
17, No. 3, May-June 2001, pp. 517-526.
"Technical Information." Technical Information | Ad Astra Rocket. Ad Astra, n.d. Web. 06 Mar
2017.

APPENDIX A: DELTA-V CALCULATIONS
Direct Lunar Transfer:
Leaving Lunar Orbit to Mars Transfer:
vΔ =
√μmoon *
2
rpark−moon
− 1
a1
−
√
μmoon
rpark−moon
with 904.9μmoon = 4 s2
km3
98600.4μearth = 3 s2
km3
32712440000.0μsun = 1 s2
km3
837.0 kmrpark−moon = 1
50295700.0 kmrearth−sun = 1
27900000.0 kmrmars−sun = 2
9.7vearth =
√
μsun
rearth−sun
= 2 s
km
189097850.0 kmatransfer = 2
r + rearth−sun mars−sun
=
2.6vreq =
√μsun * ( 2
rearth−sun
− 1
atransfer
) = 3 s
km
.9v∞−moon = vreq − vearth = 2 s
km
− 7181.1 kmalunar =
μearth
v∞−moon
= − 4
.2vlunar =
√μearth * ( 2
rmoon−earth
− 1
alunar
) = 3 s
km
.2v∞ = vlunar − vmoon−earth−orbit = 2 s
km
− 93.3 kma1 = v2
∞
μmoon
= − 9
Therefore, for leaving lunar orbit to Mars transfer:
v .6 , 00Δ =
√μmoon *
2
rpark−moon
− 1
a1
−
√
μmoon
rpark−moon
= 1 s
km
= 1 6 s
m
30

Earth Oberth Effect Transfer:
Leaving Lunar Orbit to Earth Elliptical Orbit:
vΔ =
√μmoon *
2
rpark−moon
− 1
amoon
−
√
μmoon
rpark−moon
km3
98600.4μearth = 3 s2
km3
90771.0 kmrmoon−earth = 3
.0vmoon−earth = 1 s
km
189097850.0 kmatransfer−eath = 2
r + rpark−moon moon−earth
=
.2vapogee =
√μearth * ( 2
rmoon−earth
− 1
atransfer−earth
) = 0 s
km
.8v∞,moon = vapogee − vmoon−earth = − 0 s
km
− 024.8 kmamoon = μmoon
v2
∞,moon
= − 7
Therefore, for leaving lunar orbit to earth elliptical orbit:
v .8 00Δ =
√μmoon *
2
rpark−moon
− 1
amoon
−
√
μmoon
rpark−moon
= 0 s
km
= 8 s
m
Leaving Earth Elliptical Orbit to Mars Transfer:
vΔ =
√μmoon *
2
rpark−moon
− 1
a1
−
√
μmoon
rpark−moon
with 98600.4μearth = 3 s2
km3
771.0 kmrpark−earth = 6
90771.0 kmrearth−moon = 3
0.0vearth−sun−orbit = 3 s
km
=
2.6vreq =
√μsun * ( 2
rearth−sun
− 1
atransfer
) = 3 s
km
.6v∞ = vreq − vearth−sun−orbit = 2 s
km
31

− 7976.5 kmaearth−mars =
μearth
v2
∞−earth
= − 5
202156.5 kmatransfer−earth = 2
r + rearth−moon park−earth
=
Therefore, for leaving Earth elliptical orbit to Mars transfer:
vΔ = [√μearth * ( 2
rpark−earth
− 1
aearth−mars
) −
√
μearth
rpark−earth ]
− [√μearth * ( 2
rpark−earth
− 1
atransfer−earth
)−
√
μearth
rpark−earth ]
v .404 04Δ = 0 s
km
= 4 s
m
32

For Both Direct and Earth Oberth Effect Transfer:
Launching from Surface: (Multiplying by 1.1 to account for gravity losses)
v .1Δ = 1 * √μmoon *
2
r − 100 kmpark−moon
− 1
rpark−moon
km3
v .1 .9 , 00Δ = 1 * √μmoon *
2
r − 100 kmpark−moon
− 1
rpark−moon
= 1 s
km
= 1 9 s
m
Mars Insertion:
vΔ =
√μmars *
2
rpark−mars
− 1
amars
−
√
μmars
rpark−mars
with 2828.4μmars = 4 s2
km3
576.2 kmrpark−mars = 3
4.1vmars−sun−orbit = 2 s
km
27900000.0 kmrmars−sun = 2
=
1.5vreq =
√μsun * ( 2
rmars−sun
− 1
atransfer
) = 2 s
km
.6v∞−mars = vreq − vmars−sun−orbit = − 2 s
km
− 0661.6 kmamars−earth =
μearth
v2
∞−mars
= − 6
v .5 , 00Δ =
√μmars *
2
rpark−mars
− 1
amars−earth
−
√
μmars
rpark−mars
= 1 s
km
= 1 5 s
m
Delta-v Totals:
33

Direct Lunar Transfer:
v , 00 1, 00 1, 00 975Δ = 1 9 s
m
+ 6 s
m
+ 5 s
m
= 4 s
m
Earth Oberth Effect Transfer:
v , 00 800 404 1, 00 630Δ = 1 9 s
m
+ s
m
+ s
m
+ 5 s
m
= 4 s
m
Note: Unrounded values from calculations were used for final total, hence the discrepancy in the
values listed.
Delta-v Difference:
Earth Oberth Effect Transfer is more efficient with a savings.v 45Δ = 3 s
m
34

35
COMMUNICATIONS AND DATA TEAM
TEAM LEAD: HAMILTON SUTTON
MICHAEL ABU SAADA
BROOKS JARRET
NATHAN OBHOLZ
BEN WELLS
25 APRIL 2017

36
TEAM: COMMUNICATIONS, NAVIGATION, AND DATA
SUTTON, ABU SAADA, JARRETT, OBHOLZ, WELLS
DATE: 25 APRIL 2017
OVERVIEW
The Communications, Navigation, and Data team has been tasked with fulfilling several
mission-critical functions of the Romulus Spacecraft. Effective communication between the
spacecraft and the mission control center on Earth is critical to the success of this mission. A
communications system has been selected that will have the capability of sending and receiving
all data transmissions between the spacecraft and the mission control center. Navigational
tracking functions aboard the craft have been integrated with the communications system in the
form of small packets of information that are sent and received by the communications system.
These small data packets are responsible for tracking the location of the spacecraft relative to
Earth and Mars during its transit. Finally, a robust computer and data storage system has been
selected for use aboard the Romulus spacecraft that will be responsible for managing data
storage, providing for local and ground control of the spacecraft, and allowing for data sharing
among other systems aboard the spacecraft.
TRADE STUDIES
Communications:
The antenna array will be capable of transmitting in two different radio bands, X-Band
and K-Band. X-Band radio will be used by the spacecraft for low data rate requirement tasks
such as tracking coverage, when there is significant atmospheric interference on Earth, or when a

37
direct line of sight is unavailable. The array can be switched to a higher frequency K-Band radio
signal when higher data rates are needed and a direct line of sight is possible between the
spacecraft and Earth.
The Deep Space Network, or DSN, will be used to establish constant communication
between Earth and the spacecraft. The system is composed of three large satellite dish array
compounds on Earth, each located 120 degrees latitude from each other. The location of these
stations is strategically designed to permit constant line of sight communication between Earth
and the spacecraft. The ground stations can track the spacecraft’s location as it moves away from
the Earth and so avoid coverage blackouts. As one DSN station nears the horizon due to the
rotation of the Earth and is about to drop out of sight of the spacecraft, another station becomes
visible and communications are transferred to the now visible ground station.
Radio Laser Light
Criterion: Weighting Performance Weighted
Performance
Performance Weighted
Performance
Safety 0.15 10 1.5 9 1.5
Weight 0.3 6 1.8 10 3
Lifespan 0.15 9 1.35 9 1.35
Resilience 0.1 9 0.9 9 0.9
Signal
consistency
0.3 9 2.7 4 1.2
Score: 43 8.25 41 7.8
Table 7: 10 is the highest performance or most important. Weighted performance is (weighting)*(performance).

38
When evaluating communication methods, traditional radio communications and laser
communications systems were considered. Performance values were assigned for important
characteristics of each type of system and the resulting weighted performances were determined
and totaled as shown in Table 1. Although each communications method had advantages and
disadvantages, the conventional radio communications method was shown to be more reliable
and so was selected as the primary communications system to be used for this mission.
To further its development, an experimental laser light communications system will be onboard
the spacecraft that will be tested extensively during the mission. The goal is to research the
effectiveness of laser light communications as opposed to traditional radio communications. The
laser light system will be used to send large data packages back to Earth to analyze packet loss
and overall reliability of the system. Additionally, should it prove effective, it will be used for
non-mission critical communications.
Data
The primary data storage methods studied and compared were standard Hard Disk Drive
(HDD) storage vs solid state drives (SSD), of which flash memory based solid state was chosen
for data storage on the Romulus project.
Both hard disk drives as well as solid state drives will require radiation shielding for space
applications. HDD storage drives use a physical disk to store data and have a low-cost for high-
capacity storage. The main downsides to HDD storage are susceptibility to environmental factors
such as vibration, and dramatically lower data transfer rates in comparison to flash memory.
Solid state drives have comparable storage capacities to HDD systems while also possessing
significantly faster data transfer rates, more compact systems, and greater resilience to
environmental factors. The only downside to SSD’s is their cost, which is three to four times that

39
of a HDD of similar capacity. However, this high cost does not outweigh the other benefits of
solid state drives.
Communications
The overall size of the communications system is less than fifteen cubic meters. This
value was arrived at after a visual analysis of the High-gain Antenna of the Mars Reconnaissance
Orbiter. By knowing the diameter of the dish, it was possible to calculate the protrusion depth of
the sub-reflector and the waveguide path behind the dish. These values were determined to be a
combined 1.445 meters; this value was then given a safe zone and changed to 1.5 meters while
the dish safe zone accounting for the protrusion of the Low-gain Antenna was given a larger safe
zone for operation and came to 3.5 meters. The overall size of the array would encompass 14.42
cubic meters, which was rounded to 15 cubic meters.
Using the power consumption values for the High-gain and Low-gain Antenna arrays and
varying the current input from 5-15 Amps during frequency changes, it was possible to
determine that the Voltage required for each of the active antennas would be between 6.7 and 20
Volts. In a similar manner, the amplifiers were accounted for and each determined to draw 5
Amps and 20 Volts assuming concurrent use on both the High-gain and Low-gain Antennas. The
Transponders rely on the power already supplied to the antenna and are included in its power
draw. Based on available information about the Mars Reconnaissance Orbiter, it has been
concluded that the communications system is designed such that it will only ever draw 100 Watts
at a time. The Romulus vehicle will be designed in a similar manner to prevent excessive power
draw by the communications system.

40
Transmission frequency is determined based on the type of signal that is sent and its
respective wavelength. When using X-band and Ka-band as Romulus will, the possible
transmission frequencies are respectively: 8-12 GHz, and 26-40 GHz. During mission operation,
the uplink and downlink frequencies will match that of the Mars Reconnaissance Orbiter. X-band
uplink and downlink respectively are: 7.145-7.235 GHz and 8.4-8.5 GHz. Ka-band downlink
will then be 31.8 to 32.3 GHz.
The signal strength of the transmission decays over the distance at which it is received.
Because of this, it is necessary to calculate the Free Space Path Loss (FSPL) which will give a
value based on the expected distance of the transmission. The equation below represents the
FSPL from Mars to Earth:
𝐿 𝑑𝑏,𝐹 = 92.4 + 20 log10 𝐹𝐺𝐻𝑧 𝐷 𝑘𝑚
𝐿 𝑑𝑏,𝑋 = 92.4 + 20 log10(8.4 ∗ 54.6E6) = 256.63 dB
𝐿 𝑑𝑏,𝐾𝑎 = 92.4 + 20 log10(31.8 ∗ 54.6E6) = 277.19 dB
Using the values for the frequency (8.4 GHz and 31.8 GHz) from previous calculations
and the distance from Earth to Mars (54.6 x 106
kilometers) it is possible to determine the
following FSPL values for both X-band and Ka-band respectively: 265.63 dB and 277.19 dB of
loss.
In order to determine the antenna gain, it is necessary to know the frequency, diameter of
the broadcasting dish and the aperture efficiency. The wavelength is determined from the
frequency of the signal and the speed of light. Assuming a typical aperture efficiency of 70% the
gains for X-band and Ka-band can be found:
𝐺𝑎𝑖𝑛 =
4𝜋𝐴
λ2
𝑒 𝐴 = (
𝜋𝑑
2
λ
) 𝑒 𝐴

41
𝐺 𝑑𝑏,𝑋 = 10log [𝜋(3 𝑚)2
8 ∗ 109 1
𝑠⁄
3 ∗ 108 𝑚
𝑠⁄
∗ 0.7]
2
= 56.0 dB
𝐺 𝑑𝑏,𝐾𝑎 = 10log [𝜋(3 𝑚)2
2.0 ∗ 1010 1
𝑠⁄
3 ∗ 108 𝑚
𝑠⁄
∗ 0.7]
2
64.0 dB
Thus, for the X-band frequency there will be a gain of 56 dB and for the Ka-band
frequency there will be a gain of 64 dB.
The thermal expenditures of the Communications system for the High-gain and Low-gain
antenna are each 25 Watts as they convert 25% of their energy into thermal output. The
amplifiers (assuming both are running at the same time in a worst-case scenario) both turn 15%
of their input power into heat. They have a thermal footprint of 15 Watts each. Because the
Transponder uses the existing power from the antennas its thermal footprint is included in the
High-gain and Low-gain footprint.
Data
For the data systems, the team’s main concern was that of data storage capacity, power
draw, and heat generation. To find out how much capacity was needed for the mission, the team
researched previous missions that had similar goals and duration to see what had been used in the
past. Some of the more prominent examples used were the International Space Station and the
Curiosity Mars mission. Using these as baselines the team decided to go with 12 total computers,
6 for primary usage, and 6 as backup. For each of these computers, the team wanted to bring a
high-capacity storage unit to be used in conjunction with the CPU. Once again, in reference to
previous projects, the team decided on an 8-Gbit flash NAND memory. These memory units
have a mid-range operating voltage with a long data retention life and a high number of
program/erase cycles.

42
Having a good idea of what would be running at any given time allowed the team to
calculate for power draw of the entire system.
𝑃𝑜𝑤𝑒𝑟𝑆𝑡𝑜𝑟𝑎𝑔𝑒 = 𝑉 ∗ 𝐴 = 3.6𝑉 ∗ 0.5𝐴 =
1.8𝑊
𝑢𝑛𝑖𝑡
∗ 6𝑢𝑛𝑖𝑡𝑠 = 10.8𝑊
𝑃𝑜𝑤𝑒𝑟𝐶𝑃𝑈 = 𝑉 ∗ 𝐴 = 5𝑉 ∗ 6𝐴 =
30𝑊
𝑢𝑛𝑖𝑡
∗ 6𝑢𝑛𝑖𝑡𝑠 = 180𝑊
𝑃𝑜𝑤𝑒𝑟𝑇𝑜𝑡𝑎𝑙 = 𝑃𝑜𝑤𝑒𝑟𝐶𝑃𝑈 + 𝑃𝑜𝑤𝑒𝑟𝑆𝑡𝑜𝑟𝑎𝑔𝑒 = 10.8𝑊 + 180𝑊 = 190.8𝑊
For heat generation, because there are no mechanical parts in either system, most of the
energy used is dissipated as heat energy. In order to prepare for a worst case scenario, for heat
generation calculations, it is assumed that the components function as perfect heaters and transfer
100% of their energy into heat.
𝐻𝑒𝑎𝑡 𝐺𝑒𝑛𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑀𝑎𝑥 = 190.8𝑊
𝐻𝑒𝑎𝑡 𝐺𝑒𝑛𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑆𝑡𝑎𝑛𝑑𝑎𝑟𝑑 𝑂𝑝𝑒𝑟𝑎𝑡𝑖𝑜𝑛 = 𝑀𝑎𝑥 ∗ 60% = 190.8𝑊 ∗ 0.60 = 114.48𝑊
While the heat generation characteristics of the system are important, the mass and
volume are also relevant. Using datasheets, the team calculated how much physical space the
systems would use, as well as how much they will weigh.
𝑊𝑒𝑖𝑔ℎ𝑡 𝑆𝑡𝑜𝑟𝑎𝑔𝑒 =
0.037𝑘𝑔
𝑢𝑛𝑖𝑡
∗ 12𝑢𝑛𝑖𝑡𝑠 = 0.444𝑘𝑔
𝑊𝑒𝑖𝑔ℎ𝑡 𝐶𝑃𝑈 =
2.11𝑘𝑔
𝑢𝑛𝑖𝑡
∗ 12𝑢𝑛𝑖𝑡𝑠 = 25.33𝑘𝑔
𝑊𝑒𝑖𝑔ℎ𝑡 𝑇𝑜𝑡𝑎𝑙 = 𝑊𝑒𝑖𝑔ℎ𝑡 𝐶𝑃𝑈 + 𝑊𝑒𝑖𝑔ℎ𝑡 𝑆𝑡𝑜𝑟𝑎𝑔𝑒 = 25.33𝑘𝑔 + 0.444𝑘𝑔 = 25.774𝑘𝑔
𝑉𝑜𝑙𝑢𝑚𝑒𝑆𝑡𝑜𝑟𝑎𝑔𝑒 =
3.841𝑐𝑚3
𝑢𝑛𝑖𝑡
∗ 12𝑢𝑛𝑖𝑡𝑠 = 46.1𝑐𝑚3

43
𝑉𝑜𝑙𝑢𝑚𝑒 𝐶𝑃𝑈 =
1137.04𝑐𝑚3
𝑢𝑛𝑖𝑡
∗ 12𝑢𝑛𝑖𝑡𝑠 = 13644.48𝑐𝑚3
𝑉𝑜𝑙𝑢𝑚𝑒 𝑇𝑜𝑡𝑎𝑙 = 𝑉𝑜𝑙𝑢𝑚𝑒 𝐶𝑃𝑈 + 𝑉𝑜𝑙𝑢𝑚𝑒𝑆𝑡𝑜𝑟𝑎𝑔𝑒 = 13644.48𝑐𝑚3
+ 46.1𝑐𝑚3
= 13690.58𝑐𝑚3
Communications
The communications sub-system is vital to mission success. There are two main
communication antennas aboard Romulus: a High-gain Antenna (HGA) and two Low-gain
Antennas (LGAs). The High-gain Antenna is a three-meter diameter dish antenna that is
responsible for sending data to the Deep Space Network (DSN), Earth, the Moon, or other
spacecraft at very high transfer rates. The antenna is deployed after the spacecraft has completed
the launch phase and will remain deployed and active for the duration of the mission. This
antenna will operate using a gimbal which allows it to point directly toward the receiving or
transmitting source on Earth. The total space occupied by the High-gain Antenna once deployed
will be almost fifteen cubic meters including a safe zone around the antenna.
The Low-gain Antennas have a much lower data transfer rate than the High-gain Antenna
because the radiation pattern is not focused and thus not as much of the signal will reach Earth.
However, this unfocused signal allows for communication at all times, even if the array is not
pointed towards Earth. This makes the Low-gain Antennas perfect for emergency

44
communications and orbital maneuvers. There are two Low-gain Antennas mounted to the High-
gain Antenna dish in multiple locations as detailed in Figure 1 below.
Figure 1: Rear view of the High-gain Antenna depicting both Low-gain antennas mounted to the rear and side.
In addition to the antennas, the spacecraft will have three onboard amplifiers for
transmitting. These amplifiers are high-gain, wide bandwidth and have low noise generation that
will boost the power of the antennas, these are known as Traveling Wave Tube (TWT) systems.
The amplifiers are attached to the rear of the High-gain Antenna and will ensure that
communications from the Romulus are strong enough to reach the DSN. There will be one
amplifier for X-band frequency and one for the K-band frequency as well as a backup amplifier
for the X-band frequency. The X-band frequency will then transmit using 100 watts while the K-
band will use thirty-five watts.
Transponders will be used on Romulus and will be responsible for several functions. Two
transponders will be included; the second being a backup in the event that the main transponder
has failed. Primarily, the transponders are used to translate the digital electrical signals from the
computers that are then packaged and sent in the radio broadcast. It performs an inverse
operation when receiving signals in which it translates the radio broadcast into digital electrical
signals for the onboard computers to read. The secondary function for the transponders is
automated responses. During this function the transponder passively listens for specific signals
and replies automatically based on the message received. The transponders are crucial for the

45
navigation system as they quickly and efficiently transmit important navigation data to Earth
which allows for analysis and location determination for the spacecraft (See Tracking Coverage
below).
Tracking Coverage:
The Romulus spacecraft’s positional determination function, or “Tracking
Coverage,” will be accomplished through the use of radio signals sent and received between the
communications system onboard the spacecraft and the Deep Space Network located on Earth.
To determine the location of the spacecraft at any point in time, a location data packet will be
sent from the DSN to Romulus. Upon receiving this data packet, the onboard Navigation system
identifies it as a location request and immediately retransmits the same signal back to Earth.
When the return signal reaches the DSN, the distance to the Romulus is calculated. This
calculation is performed by taking the amount of time between the initial sending of the data
packet and the return of the data packet from the ship, subtracting the turnaround time, dividing
the result by two, and then multiplying times the speed of light. This process yields a distance
measurement accurate to about ten meters.
Further location determinations will be performed using a method called Delta
Differential One-Way Range, or Delta DOR. Delta DOR uses two widely separated antennas on
Earth to simultaneously track the data transmission delays between the spacecraft and Earth. By
comparing lag times between the two stations the relative angle between Earth and the spacecraft
can be determined (Figure 2). Time tracking of this angular measurement between the Romulus
and the two antennas can then be used to calculate the speed and location of the spacecraft in the
lateral direction. To further enhance the accuracy of this method, both ground station antennas
also lock onto the naturally occurring electromagnetic radiation signals emitted by a quasar.

46
Provided that the location of the quasar is already known and is at a similar angle relative to
Earth as to that of the spacecraft, the signal disturbances due to Earth’s atmosphere can be
determined. The error can be reduced and the relative angle calculation can be corrected to be
accurate to within five to ten nano-radians.
Figure 2: A diagram of the Delta-DOR process.
The final navigational locating method to be used aboard the Romulus spacecraft is
Doppler data. When the comms array on the Romulus sends a location data packet back to Earth,
the frequency of the radio wave experiences a Doppler shift due to the velocity of the spacecraft
relative to Earth. When the ship is moving away from Earth, the emitted frequency will shift to
be lower, and when the ship is moving towards Earth, the frequency shift will be higher. As the
true frequency of the emitted radio signals is known, the ground station can then determine from
the Doppler shift how fast the ship is moving in the lineal direction from Earth.
Data
The Data system will take in data from the entire spacecraft and all the modules and
experiments as well as all subsystem data. It will store and catalogue all relevant data from
modules and send any urgent or necessary data to the Communications array in order to be sent
to the appropriate Command Center and processed. It will encrypt all data that is output from the

47
system using at minimum AES128 standards. The system will also organize subsystem and
module data so that the onboard crew can access and process it via terminals aboard the
spacecraft.
The ideal computers for the Romulus mission are those in the RAD 5500 series. The 5500
series is the successor to the RAD 750, which has been heavily used in space applications
including the Juno mission, mars and lunar missions, and lastly satellite missions. Specifically,
the RAD 5515 is a radiation hardened computer designed by BAE systems for spacecraft.
Notable specs for the RAD 5515 include:
 Maximum Power draw and dissipation: 13.7W
 Operating Temperature: -55 to +125 ℃
 Voltages:
 Core: .95 V
 I/O: 1.8. 2.5, 3.3 V
 Radiation: 1 MRad maximum dosage
 Memory: 64 Gb
 Processor throughput: 1.4 GOPS
As for data storage, the CI-1HN89 8-Gbit Rad tolerant NAND flash memory by
Telecommunications systems has been selected for mass storage purposes.

48
Figure 3: An image of the Telecommunications systems CI-1HN89 NAND device.
These flash memory storage units are radiation hardened and have been used in
spacecraft before. They feature non-volatile memory and very low power solid state data
recording. The memory unit chosen features 100,000 program/erase cycles, 10-year data
retention lifespan, operating voltage of 2.7-3.6 volts, and resilient radiation characteristics (See
appendix A) all while remaining rather compact.
The Communications, Navigation, and Data systems aboard the Romulus spacecraft are
required to perform several complex functions, often simultaneously, and must be reliable
enough to guarantee mission success. The data system selected makes use of the most reliable,
and functional space grade computing and storage hardware available for mission applications.
The communications system utilizes highly proven equipment, extensively tested in both short
and long range space missions. Finally, the navigation system will make use of multiple proven
positional location methods that utilizing the communications and data hardware already
onboard. Each system in the design has several redundancies for fault tolerance, while easily
fulfilling the functions required of them. This ensures that the neither the success of the mission
nor the lives of the crew are ever in any danger due to a failure from these vital systems.

49
REFERENCES
"About Delta DOR." European Space Agency. N.p., 1 Oct. 2013. Web. 17 Apr. 2017.
“Antenna Fundamentals.” National Radio Astronomy Observatory. Web. 15 Mar. 2017.
“Comtech 8-Gbit Rad Tolerant NAD Flash.” Comtech Telecommunications. Web. 12 Apr. 2017
“Mars Reconnaissance Orbiter: Antennas.” NASA. NASA. Web. 13 Mar. 2017.
"Navigation - Mars Science Laboratory." NASA. NASA, n.d. Web. 17 Apr. 2017.
“Space Communications with Mars.” Luxorion. Web. 24 Mar. 2017.
“Stevens, David P. "Flight to Mars: How Long? Along What Path?" Educational Web Sites on
Astronomy, Physics, Spaceflight and the Earth's Magnetism. N.p., 12 Dec. 2004. Web. 24
Apr. 2017.
“Transmitting and Receiving Antennas.” Rutgers University. Web. 15 Mar. 2017.
“Wilson, J. "Homann Transfers." The University of Georgia. N.p., n.d. Web. 24 Apr. 2017.

50
APPENDIX A: RADIATION CHARACTERISTICS
OF CI-1HN89 NAND FLASH MEMORY (FROM DATASHEET)
Radiation Characteristics:
 TID: 50krad(Si) High Dose Rate (MIL-STD-883, TM1019 Cond.A)
 SEL immune to LETs ≤ 75.6 MeV-cm2/mg; at Ta = +85°C
 No destructive SEFI ≤ 75.6 MeV-cm2/mg, under unbias, static bias and read mode
 No destructive SEFI ≤ 50.9 MeV-cm2/mg, under erase/write mode 1
 SEU threshold: ~3 MeV-cm2/mg
 SEU saturated cross-section is ~3.5E-11cm2/bit 2/

51
ELECTRICAL POWER TEAM
TEAM LEAD: GARRET SENTI
THOMAS LEPOIDEVIN
DANIEL WENGER
MIKE ZINK
25 APRIL 2017

52
TEAM: ELECTRICAL POWER
SENTI, LEPOIDEVIN, WENGER, ZINK
DATE: 25 APRIL 2017
OVERVIEW
The Electrical Power team of the Romulus Mission is responsible for selecting the power
generation, storage, and distribution systems for the spacecraft. Through research and trade
studies the team determined what equipment should be used as the major parts of the electrical
power system. Once the specific equipment was selected the team prepared information to give
to the other subsystems about how the size and weight of the parts of the system would have to
increase to accommodate increasing electrical loads. With this information in mind the other
subsystems then gave the Electrical Power team estimates for the electrical power that their
systems would require. Using these estimates the team was then able to finalize estimates for the
volume and mass for the major components of the electrical power system that would be needed
for the Romulus Mars Mission.
TRADE STUDIES
After presenting to the class and receiving feedback from the overview team, the
electrical team performed tradeoff studies for the different options that were considered for
power collection and power storage (Appendix A).
The first tradeoff study compared the efficiency of the power generation devices to its
relative max safety, size and mass, lifespan of the system, and cost (Appendix A.1). Since the
team was more concerned about the safety of the crew, it became clear that the solar panels were

53
the best option, even though they take up a large amount of space and mass. The solar panels
also proved to have a lower cost and longer lifespan than both the radioisotope thermal generator
and nuclear fission processes. It was clear that the solar panels would be the most effective
choice for the mission.
The second tradeoff study determined which power storage method would be used. The
categories were safety, energy density, the battery’s ability to hold a charge, and consistency in
performance (Appendix A.2). Since there is no possibility for an emergency evacuation in flight,
safety was a heavy concern. The PEM hydrogen fuel cell results show that there is little risk to
the crew since the only waste the fuel cells will produce is water and oxygen. However, in
regards to energy density and the ability to hold a charge, the Lithium Ion battery is the best
option. From this trade off study, the Lithium Ion battery was shown to be the most favorable
option based on the conditions established.
The first calculations performed were to determine the size of the solar panels needed to
power the spacecraft. The team assumed that the spacecraft would be using solar panels like that
of the Juno Satellite, in which the length and width of one of the arrays are 8.2 meters and 2.7
meters, respectively, meaning that the total area is 72 meters squared, since there is a total of
three arrays. From the test performed, the solar panels produced 12 to 14 kilowatts of power on
the Earth. Using the equation:
𝐴𝑣𝑒𝑟𝑎𝑔𝑒(12000,14000)𝑊
72𝑚2
= 180
𝑊
𝑚2
the value calculated helped the team determine the size the solar panels need to be to generate
the required power for the spacecraft. The team knew that since solar panels are being used for
the mission and the spacecraft will be traveling farther away from the sun then that would mean

54
that the size of the solar panels would have to be larger to account for that change. Research was
conducted to determine the changes in wattage produced at different locations in space. The
value that the team determined that the power for the entire system would need to be divided by
is 0.43 so that the amount of power generated from Earth’s orbit would decrease once it reaches
Mars but still can satisfy the conditions required (Appendix B). The power needed for the overall
system would be 16.7 kilowatts but the team include a safety factor of 1.25 making the desired
power of generation 21 kilowatts which means that from the solar panels would be designed to
generate 49 kilowatts in Earth’s orbit. This means that the total area of the solar panels would be
270 meters squared (Appendix C). Since the solar panels chosen by the team are based from
Juno’s solar panels, which have a total mass of 340 kilograms to generate an average of 13
kilowatts of power, the mass of the solar panels needed to produce 49 kilowatts at Earth’s orbit
would weigh around 1400 kilograms (Appendix D). The reason for using the solar panel type
designed for Juno was the high TRL level associated with its use, along with its outstanding
performance on Juno.
The team performed battery calculations to determine the number of lithium-ion cells that
would be required to operate the spacecraft for 10 hours in case of emergency, eclipse, or
technical issues. Assuming the spacecraft draws 21 kilowatts of power for those 10 hours, the
total stored electrical energy needs to be 210 kilowatt hours. Using GS Yuasa LSE190 lithium-
ion cells, which are related to those currently used on the ISS and each have a capacity of 758
Watt hours (Appendix E), the team determined that 280 cells would be needed. To calculate the
heat generation from the batteries during discharge, the team started with battery heat generation
equations from William Walker’s “Short Course on Lithium-Ion Batteries” and then modified

55
the results to account for extremely slow (10-hour) discharge rate. The team determined that the
heat generation from the battery system may be about 0.7 kilowatts during discharge.
The Power Control Distribution Unit (PCDU) is the central unit of circulating the watts
that are generated or stored to be circulated towards all the components that require power. The
team gathered most of their information from Thales Alenia Space Company which provided
basic information on what to look for in a PCDU as well allowing the team to figure on that each
unit should be able to distribute 6.5 kilowatts which means that four PCDUs should be used
since the entire spacecraft would only need to obtain 21 kilowatts. The dimensions and the mass
of a single PCDU was determined by making an educated assumption based on information
provided by Thales Alenia Space and Therma Space which allowed the team to determine that
the length, width, and height of one unit would be 600 millimeters, 345 millimeters, and 195
millimeters, respectively as well as saying that one unit would have a mass of 30 kilograms.
Since the team could not find enough information could be found on the heat generated, the team
assumed that each unit would produce 400 watts of heat.
Since the subsystems vary in power requirements throughout the craft, the cables used in
the electrical system would vary based on the power needed to be transferred across the system.
Since the cable size used in a spacecraft varies between the sizes of #1/0 AWG to #30 AWG and
it is difficult to determine at this point how much of each one would be used, the team assumed
that the #16 AWG would represent all the wire in the spacecraft. Based on the design of the
spacecraft, the team decided that a 100 miles of wire would be sufficient. Using the chart
provided by Colonial Wire, the team converted the 100 miles to feet to them find the mass of the
wires based upon the information provided which lead to the team to calculate a value of 1950
kilograms for the mass of all the wires used on the spacecraft. The team then found the diameter

56
of the #16 AWG wire which the team used the minimal size for the diameter which was 1.5
millimeters and then using the equation:
𝑉 = 𝜋 ∗ 𝑟^2 ∗ 𝑙
determined that the total volume would be 0.28 meters cubed. The team also determined that the
amount of heat that would be generated would be around 5 kilowatts. Although it is hard to
determine the exact values that would be used in spacecraft, the team believes this does set an
idea of what to expect.
The team knows that there would be more components that would be used in the real
design but since it is impossible to determine what would be used, the team focused on the
essentials for the mission which were power generation, power storage, transfer, and regulations.
Overall, the team has established a foundation to work from and what could be expected with the
finalized design of the entire spacecraft.
Solar Panels
The team concluded that this mission will require solar panels with a total surface area of
approximately 270 square meters to generate sufficient power for the Romulus spacecraft during
its entire mission to Mars. This area will be divided into two solar arrays, one on either side of
the spacecraft. Each array will be 27 meters long and 5 meters wide. The solar panels have a
thickness of 50 millimeters, which means the total minimum volume of the solar panels will be
13.5 cubic meters, or 6.75 cubic meters per array, see Figure 1 for an illustration of what the
design of the panel would potentially look like. The weight of the solar panels will be a total of
1400 kilograms. At the start of the mission, the solar panels will generate around 49 kilowatts of
power and once the spacecraft reaches Mars then the amount of power that it would be able to

57
generate is 21 kilowatts. In terms of heat load, the solar panels will produce a total of 10
kilowatts due to the energy absorbed but not converted to electricity.
o Total Number of Wings: 2
o Total Area: 270 m2
o Total Volume: 13.5 m3
o Total Mass: 1400 kg
o Power Produced at Start of Mission: 49 kW
o Power Produced at End of Mission: 21 kW
o Total Heat Load: 10 kW
Figure 1: A Solar Panel Wing in Space
Li-Ion Batteries
To provide 210 kWh of power storage, the Romulus mission spacecraft will be supplied
with a total of 280 LSE190 lithium-ion cells (Figure 2, battery on right). These cells will be
grouped into 7 battery units, each 3.2 meters long and 0.35 meters wide. Altogether, these will
take up a volume of 2.9 cubic meters and have a mass of 1285 kilograms. When the batteries are

58
being charged, or used to power the spacecraft, they may generate about 0.7 kilowatts of heat
energy. The heat generated from charging the batteries will not exceed the maximum heat
generated from discharging them if the charging rate is not allowed to exceed the discharge rate.
o Total Number of Batteries: 7
o Total Number of Cells: 280
o Total Heat Load: 0.7 kW
Figure 2. GS Yuasa High-Capacity Batteries for Space Operations
Power Control Distribution Unit (PCDU)
These devices will enable power to flow from the batteries to the instruments that require
power. For the distribution of the power, the team assumed there will be around 4 different

59
power distribution units in which the total volume of all the PCDU would be 0.162 meters cubed
with a total mass of 120 kilograms, see Figure 3 for what a single PCDU would look like. The
heat load of all the PCDU is assumed to be 1.6 kilowatts. Since there are not many suppliers of
PCDUs capable of the high-output necessary for deep-space travel, these calculations were based
from the use of PCDU systems from Surrey Satellite Technologies LTD, and English company
that creates space-worthy electronics.
o Total Number of PCDU: 4
o Total Heat Load: 1.6 kW
Figure 3: Representation of the PCDU for the Spacecraft
Power Cables
The power needed to be transferred throughout the entire spacecraft is around 21
kilowatts. To transfer this around the craft to the other systems this electricity needs wires to go
through. Calculations for the wires assume that all wires used be using the same gauge of #16

60
AWG (Figure 4). Even though in actuality there would be other wire sizes used, for the sake of
estimation this average size was used. It was assumed that around 100 miles of wire would be
used and these wires would have a diameter of 1.5 millimeters. With these assumptions, the total
amount of power cables takes up 0.28 meters cubed and weighs 1950 kilograms. These wires
generate an estimated 5 kilowatts of heat energy when being used.
o Total Heat Load: 5 kW
Figure 4: #16 Gauge Electrical Wire
The Electrical Power team was charged with the determining methods and application of
power generation, power storage, and power distribution systems for the spacecraft. Through
research, tradeoff studies, calculations, estimates, and communication with the other subsystem
teams, the team could create an outline of the major components of the electrical power system
that would be required for the Romulus Mission. To generate power the team concluded that
solar panels should be used. For power storage, Lithium Ion batteries were chosen. Finally, to
distribute this power about the craft, the team chose to use Power Control Distribution Units, and

61
power cables. These components were found to best meet the demands for the mission compared
to other options considered. The outline of the system components specifies the number of each
component, as well as their volume and mass, in the electrical power system that are estimated to
be needed for the mission. The estimated heat load of each major part was included to assist the
Thermal Control Team develop their system by giving more accurate estimates for them to use.
This outline shows how the Electrical Power team would plan to meet the requirements of such a
mission as the Romulus Mission to Mars.

62
REFERENCES
"Ampacity Charts." Cerrowire. Cerrowire, 2014. Web. 17 Apr. 2017.
Axon Cable. "Cables & Harnesses for Space Applications." Cables & Harnesses for Space
Applications - Axon Cable. Axon Cable, Oct. 2014. Web. 23 Apr. 2017.
Colonial Wire. "Wire Weights Per 1000 Feet (in Pounds)." Wire Weights per 1000 Feet (in
Pounds) - Colonial Wire. Colonial Wire & Cable CO. Of New Jersey, INC., n.d. Web. 18
Apr. 2017.
Dismukes, Kim, and Amiko Kauderer. "The 21st Century Space Shuttle." The Most Complex
Machine Ever Built - NASA Human Space Flight. NASA, 20 Jan. 2010. Web. 23 Apr.
2017.
Dunbar, Brian. "Juno Solar Panels Complete Testing." NASA. NASA, 24 June 2016. Web. 17
Apr. 2017.
Dunbar, Brian. "Juno's Solar Cells Ready to Light Up Jupiter Mission." NASA. NASA, 27 June
2016. Web. 17 Apr. 2017.
Gaston, Darilyn M. "Selection of Wires and Circuit Protective Devices for STS Orbiter Vehicle
Payload Electrical Circuits." NASA Technical Memorandum 102179. Lyndon B.
Johnson Space Center, n.d. Web. 23 Apr. 2017.
McClure, Bruce. "What Is an Astronomical Unit?" What Is an Astronomical Unit? | Space |
EarthSky. EarthSky, 21 Oct. 2016. Web. 22 Apr. 2017.
Panduit. "Electrical Wire Sizes Selection Guide." WW-WASG03 Electrical Wire Sizes.qxp -
Panduit. Panduit, July 2011. Web. 18 Apr. 2017.

63
Saito, Yoshiyasu, Masahiro Shikano, and Hironori Kobayashi. "Heat Generation Behavior
during Charging and Discharging of Lithium-ion Batteries after Long-time Storage."
Journal of Power Sources 224 (2013): 294-99. ScienceDirect. Web. 24 Apr. 2017.
Thales Alenia Space. "Power Conditioning and Distribution Unit Medium Power." PCDU
Medium Power - Thales. Thales Alenia Space, Feb. 2014. Web. 18 Apr. 2017.
Therma Space. "Power Conditioning & Distribution Unit." Power Conditioning & Distribution
Unit - Terma. Therma Space, n.d. Web. 18 Apr. 2017.
Walker, William. “Short Course on Lithium-Ion Batteries.” NASA. NASA, 2015. 10 Apr. 2017
Wudka, Jose. "The Inverse-Square Law." The Inverse-Square Law - UCR Physics. UCR Physics,
24 Sept. 1998. Web. 23 Apr. 2017.

64
Appendix A: TRADE-OFF STUDIES
Appendix A.1: Power Source Trade-Off Study Results

65
Appendix A.2: Power Storage Trade-Off Study Results

66
APPENDIX B: SOLAR ENERGY FLUX
Solar Energy Flux Related and Distance from Sun for All Planets
Comparison Between the Earth and Mars

67
Solar Energy Flux Compared to Distance from Sun
Planet Distance from Sun (AU) Ratio of Energy Produced (Earth = 1)
Earth 1 1
Mars 1.524 0.43
Jupiter 5.203 0.04
Saturn 9.54 0.01
Uranus 19.18 0.003
Neptune 30.06 0.001
Pluto 39.53 0.0006
Table of Values used for Graph

68
APPENDIX C: POWER VS SOLAR PANEL AREA
Solar Panel Power Generated to Total Area Required
This is assuming 180 W/m2
for the Solar Panels

69
APPENDIX D: POWER VS SOLAR PANEL MASS
Solar Panel Power Generated to Total Mass of the Solar Panels
This is assuming 36 W/kg for the Solar Panels

70
APPENDIX E: BATTERIES
GS Yuasa LSE190 Specifications

71
Appendix E
GS Yuasa LSE190 Specifications

72
THERMAL CONTROL TEAM
TEAM LEAD: JONATHAN AUKES
ANDREW JOHNSON
NICK MALINARIC
SAM SMITH
25 APRIL 2017

73
TEAM: THERMAL CONTROL TEAM
AUKES, JOHNSON, MALINARIC, SMITH
DATE: 25 APRIL 2017
OVERVIEW
The thermal control team was tasked with collecting and expelling heat generated
throughout the spacecraft. Spacecraft subsystems, components, and crew all possess safe
operating temperature ranges and limits. The thermal control system will maintain all
subsystems, components, and crew within their respective temperature ranges. To do so, a
combination of exterior insulation, heat generation, heat transfer, and heat rejection solutions
will be employed. The specific solutions to these general areas of focus are detailed below.
TRADE STUDIES
Weighted evaluations were completed for each of the systems within the Thermal Control
System to determine which option best met the needs of each system.
Insulation
Coatings MLI Foam
Criterion: Weighting Performance
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Conduction Resistance 0.2 1 0.2 9 1.8 10 2
Radiative Resistance 0.4 8 3.2 8 3.2 2 0.8
Lifespan 0.2 4 0.8 8 1.6 4 0.8
Weight 0.2 9 1.8 8 1.6 4 0.8
Score: 22 6 33 8.2 20 4.4
Table 1: Insulation Weighted Evaluation

74
The weighted evaluation of the three options for spacecraft insulation (Table 1) showed
that Multi-Layer Insulation (MLI) was indeed the best option. MLI involves Kapton in layers
with polyethylene terephthalate, polyester, Mylar, or Teflon, providing adequate emissivity and
absorptivity for the vessel. This option is preferred for its well-proven reputation, adaptability to
structural and thermal system design, and integration capability with other subsystems.
Heat Generation
Foil Heater RHU
Criterion: Weighting
Performan
ce
Weighted
Performanc
e
Performan
ce
Weighted
Performanc
e
Efficiency 0.3 5 1.5 9 2.7
Controllability 0.3 10 3 2 0.6
Lifespan 0.1 7 0.7 5 0.5
Weight 0.3 8 2.4 5 1.5
Score: 30 7.6 21 5.3
Table 2: Heat Generation Weighted Evaluation
The weighted evaluation, shown in table 2, confirmed that out of the two options for heat
generation, the Electric foil heaters were the better option. Electrical Foil Heaters involve
resistive wires in Kapton or adhesive becoming hot when a current is passed through them,
conducting that thermal energy to the cooler surroundings. This option is recommended due to
its low cost, versatility, and easy automation adaptability.

75
Electrical System Cooling
Peltier Vapor-Compression Liquid Cooling
Criterion:
Weightin
g
Performance
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Efficiency 0.4 7 2.8 10 4 9 3.6
Simplicity 0.2 10 2 1 0.2 6 1.2
Transfer
Rate 0.3 1 0.3 8 2.4 6 1.8
Weight 0.1 3 0.3 5 0.5 7 0.7
Score: 21 5.4 24 7.1 28 7.3
Table 3: Electrical System Cooling Weighted Evaluation
While the Vapor-Compression scored well in the weighted evaluation (Table 3), it still
came in second to the Liquid Cooling option. Liquid Cooling involves piping a refrigerant fluid
around the circuitry, absorbing heat into the pipe through conduction and into the fluid by
convection, to meet the requirement for expelling thermal energy from the electrical systems.
This method is preferred for its balance between effectiveness and proven reliability.
Heat Transfer
Mechanical CCHP VCHP Fluid Loop HTPL
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Operation 0.2 5 1 5 1 6 1.2 9 1.8 9 1.8
Simplicity 0.3 10 3 6 1.8 4 1.2 8 2.4 6 1.8
Transfer Rate 0.4 1 0.4 6 2.4 7 2.8 7 2.8 9 3.6
Weight 0.1 1 0.1 6 0.6 6 0.6 7 0.7 7 0.7
Score: 17 4.5 23 5.8 23 5.8 31 7.7 31 7.9
Table 4: Heat Transfer Weighted Evaluation
The weighted evaluation (Table 4) with the selected criterion and weighs showed that
Hybrid Loops (HTPL) were the best option to go along with the selected liquid electronic system

76
cooling. Hybrid Loops involve circulating a diphasic fluid through capillary action with the
addition of mechanical pumping if necessary, thus maintaining the craft’s temperature balance
and fulfilling the requirement for expelling thermal energy from other spacecraft subsystems.
This option seems most promising due to its high reliability, relative efficiency, and volume of
heat transfer.
Heat Transfer Coolant
Water Ammonia Triol Polymethyl Siloxane Freon-218
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Performance
Weighted
Performance
Safety 0.2 10 2 3 0.6 9 1.8 10 2 8 1.6
Operating Range 0.3 3 0.9 6 1.8 5 1.5 5 1.5 7 2.1
Transfer Rate 0.4 8 3.2 9 3.6 9 3.6 5 2 5 2
Weight 0.1 4 0.4 6 0.6 4 0.4 4 0.4 5 0.5
Score: 25 6.5 24 6.6 27 7.3 24 5.9 25 6.2
Table 5: Heat Transfer Coolant Weighted Evaluation
Due to the differing needs of interior and exterior heat transfer, the team chose to select
the two highest scoring options from the weighted evaluation, shown in table 5. Ammonia and
Triol have very high heat capacity and low freezing points, allowing for a greater flexibility than
water in cooling systems. Triol is recommended for internal cooling due to its low toxicity and
easy detectability, while ammonia is recommended for external cooling since toxicity is
mitigated in this environment. Both these solutions are well proven and meet the requirement of
transporting thermal energy throughout the spacecraft.

77
Heat Rejection
Fixed Radiator Variable Radiator
Performan
ce
Weighted
Performanc
e
Performan
ce
Weighted
Performanc
e
Efficiency 0.2 6 1.2 9 1.8
Adjustability 0.3 1 0.3 6 1.8
Simplicity 0.2 10 2 8 1.6
Weight 0.3 8 2.4 7 2.1
Score: 25 5.9 30 7.3
Table 6: Heat Rejection Weighted Evaluation
The weighted evaluation, shown in table 6, confirmed that the better option for heat
rejection was the Variable Emissivity Radiation system. Variable Emissivity Radiation allows
for variations in surface emissivity through mechanical louvres, lessening the need for additional
heat generation should the vessel become too cold and fulfilling the requirement to maintain
appropriate operating temperatures. This option seems best since it adds little extra complexity
and provides an additional method to prevent overcooling and lower energy consumption.

78
Cabin Thermal Control
Heat Exchangers Electronic Systems
Performan
ce
Weighted
Performanc
e
Performan
ce
Weighted
Performanc
e
Efficiency 0.4 9 3.6 3 1.2
Adjustability 0.2 4 0.8 6 1.2
Reliability 0.2 7 1.4 5 1
Weight 0.2 7 1.4 5 1
Score: 27 7.2 19 4.4
Table 7: Cabin Thermal Control Weighted Evaluation
Based off the selected criteria and weights, the weighted evaluation (Table 7) showed the
team that the better option for controlling the cabin temperature was the Heat Exchanger option.
Heat exchangers would be combined with a variety of temperature sensors and fluid valves to
either run cool or hot water throw the heat exchangers to maintain a comfortable cabin
temperature for the crew. This option is the best option because of its efficiency as it recycles the
heat gathered from heat emitting electrical components in the spacecraft.
Calculations were performed for each of the systems for evaluate the electrical power
consumed and thermal power produced with respect to the mass of the system and amount of
thermal power that the system could reject or move. The numbers were ultimately combined
with thermal loads from the other subsystems in the spacecraft to calculation the total mass of the
Thermal Control System.

79
Insulation
Insulation is key to controlling heat flux in the spacecraft. MLI will be the insulation type
investigated, as it is plays the primary insulative role aboard the spacecraft. MLI consists of a
varying number of layers of radioactively reflective foil separated by thin netted layers. The
number of layers per centimeter and the total thickness of material are the main variables.
Though the radiation term dominates, all three main heat transfer modes are present in the
empirical and general physics formulas in the following equations. For a given thickness, there is
a balance between the number of layers of reflective foil and the total heat flux through the
medium. Too many layers produce a larger conduction coefficient through the tightly-packed
spacing material. Therefore, there exists an optimal number of layers of foil and spacing material
for a given total thickness. This given thickness is often a function of the allowable mass of the
insulation. The following Figure 1 shows a number of empirical based functions modeling this
phenomena. Figure 1 shows the effect of changing the number of reflective layers while a
thickness of 25.4 mm is fixed.

80
Figure 1: Heat Flux vs Layer Density for 25.4mm Thickness
Continuing to the find the total mass of an insulation system from the optimal layer
thickness and density is not straightforward and NASA has entire FORTRAN programs designed
to do such calculations. For preliminary design, a 40 shield layer design with a 20 layers/cm
density will be used. The two most important numbers at this point in the design phase are
density per unit area and the heat flux. Using the following Figure 2, which holds true for a
constant 40 shield layers, yields a rough surface area density of 0.3875 kg/m^2. This case would
also result in a total heat flux from radiation of 0.18 W/m^2. This allowed the team to find total
heat transfer rate through the spacecraft after the exterior dimensions and surface area were
calculated.

Romulus Project Mission Report from the Moon to Mars

Romulus Project Mission Report from the Moon to Mars

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