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AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - i
Table of Contents
1. Project Vision Overview ............................................................................................................1
1.1. Foreword by Professor James M. Longuski .......................................................................2
1.2. Acknowledgments ..............................................................................................................4
1.3. Project Team .......................................................................................................................5
1.4. Vision Statement .................................................................................................................6
2. Project Introduction ...................................................................................................................7
2.1. Report Organization ............................................................................................................8
2.2. Project Objective ................................................................................................................9
2.3. Mission Design Requirements ..........................................................................................10
2.4. Design Process ..................................................................................................................13
3. Mission Overview and Timeline ..............................................................................................21
3.1. Quick Reference for Vehicle Specifications .....................................................................22
3.2. Vehicle Overviews ............................................................................................................31
3.3. Scientific Overview ..........................................................................................................62
4. Project Conclusions .................................................................................................................67
4.1. General Project Concerns .................................................................................................68
4.2. Detailed Mission Timeline ...............................................................................................80
4.3. Estimated Mission Cost ....................................................................................................84
4.4. Risk Assessment ...............................................................................................................95
4.5. Closing Comments..........................................................................................................100
5. Detailed Vehicle Descriptions ...............................................................................................101
5.1. Supply Launch Vehicle ...................................................................................................102
AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - ii
5.1.1. Launch Vehicle Selection ..............................................................................102
5.1.2. Launch Manifest and Timeline ......................................................................104
5.2. Supply Transfer Vehicle .................................................................................................109
5.2.1. Construction in LEO ......................................................................................109
5.2.2. Configuration Overview ................................................................................112
5.2.3. Trajectory ......................................................................................................113
5.2.4. Power Systems ...............................................................................................119
5.2.5. Propulsion Systems .......................................................................................136
5.2.6. Attitude Determination and Control Systems (ADCS) .................................142
5.2.7. Structures and Thermal Systems ...................................................................148
5.2.8. Communications Systems .............................................................................156
5.2.9. Ceres Operations ...........................................................................................162
5.2.10. End of Life Configuration ...........................................................................165
5.3. Crew Launch Vehicle .....................................................................................................167
5.3.1. Launching the Crew ......................................................................................167
5.3.2. Crew Launch Manifest and Timeline ............................................................168
5.4. Crew Transfer Vehicle ....................................................................................................173
5.4.1. Construction in LEO ......................................................................................173
5.4.2. Configuration Overview ................................................................................177
5.4.3. Outbound Trajectory .....................................................................................181
5.4.4. Power Systems ...............................................................................................187
5.4.5. Propulsion Systems .......................................................................................196
5.4.6. Attitude Determination and Control Systems (ADCS) .................................216
AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - iii
5.4.7. Human Factors Systems and Habitability Concerns .....................................240
5.4.8. Structural Systems .........................................................................................285
5.4.9. Thermal Control Systems ..............................................................................289
5.4.10. Aerodynamic Systems .................................................................................291
5.4.11. Communications Systems ...........................................................................304
5.4.12. Rendezvous with Crew Capsule ..................................................................310
5.4.13. Ceres Operations .........................................................................................311
5.4.14. Return Trajectory .........................................................................................317
5.4.15. Aerodynamic Maneuvers .............................................................................321
5.4.16. End of Life Configuration ...........................................................................326
5.5. Crew Capsule ..................................................................................................................329
5.5.1. Configuration Overview ................................................................................329
5.5.2. Power Systems ...............................................................................................331
5.5.3. Propulsion Systems .......................................................................................334
5.5.4. Human Factors Systems and Habitability Concerns .....................................337
5.5.5. Attitude Determination and Control Systems (ADCS) .................................340
5.5.6. Structural and Thermal Systems ....................................................................342
5.5.7. Aerodynamic Systems ...................................................................................349
5.5.8. Communications Systems .............................................................................362
5.5.9. Crew Capsule Operations ..............................................................................364
5.5.10. Storage and Return of Ceres Rocks .............................................................374
5.5.11. Aerodynamic Maneuvers .............................................................................375
5.6. In-Situ Propellant Production Stations ...........................................................................381
AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - iv
5.6.1. Configuration Overview ................................................................................381
5.6.2. ISPP Production Timeline .............................................................................384
5.6.3. Power Systems ...............................................................................................386
5.6.4. Harvesters Detailed Description ....................................................................398
5.6.5. Extractor Detailed Description ......................................................................406
5.6.6. Tank and Vehicle Connections ......................................................................416
5.6.7. Communication Systems ...............................................................................417
5.6.8. End of Life Configuration .............................................................................418
5.7. Exploration Rovers .........................................................................................................419
5.7.1. Configuration Overview ................................................................................419
5.7.2. Power Systems ...............................................................................................421
5.7.3. Propulsion Systems .......................................................................................426
5.7.4. Human Factors Systems and Habitability Concerns .....................................433
5.7.5. Attitude Determination and Control Systems ................................................435
5.7.6. Structural and Thermal Control Systems ......................................................439
5.7.7. Communication Systems ...............................................................................454
5.7.8. Ceres Rock Collection Process ......................................................................457
5.7.9. Science Toolbox and Experimentation ..........................................................458
5.7.10. Autonomous Operations ..............................................................................459
5.7.11. End of Life Configuration ...........................................................................463
5.8. Rescue Rover ..................................................................................................................465
5.8.1. Configuration Overview ................................................................................465
5.8.2. Power Systems ...............................................................................................466
AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - v
5.8.3. Propulsion Systems .......................................................................................470
5.8.4. Human Factors Systems and Habitability Concerns .....................................477
5.8.5. Attitude Determination and Control Systems ................................................485
5.8.6. Structural and Thermal Control Systems ......................................................439
5.8.7. Communication Systems ...............................................................................499
5.8.8. Trajectory and Flight Path .............................................................................501
5.8.9. Autonomous Operations ................................................................................503
5.8.10. End of Life Configuration ...........................................................................505
5.9. Communications Network ..............................................................................................507
5.9.1. Ceres Orbiting Satellites ................................................................................507
5.9.2. Earth-Trailing Relay Satellite ........................................................................534
5.9.3. Earth-Orbiting Relay Satellite .......................................................................552
AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - vi
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Project Overview Page 1
1. Project Overview
“When Earth was tamed and tranquil, and perhaps a little tired, there would still be scope for
those who loved freedom, for the tough pioneers, the restless adventurers. But their tools would
not be ax and gun and canoe and wagon; they would be nuclear power plant and plasma drive
and hydroponic farm. The time was fast approaching when Earth, like all mothers, must say
farewell to her children.”
- Arthur C. Clarke, 2001: A Space Odyssey
Project Overview Foreword Page 2
Author: Prof. James Longuski
1.1. Foreword
This report represents the culmination of an intensive spacecraft design course, AAE 450,
undertaken by seniors during a single semester. The students perform a feasibility study for a
specified mission goal, subject to certain constraints.
The entire class works as a single team to achieve this goal. They elect a Project Manager and
an Assistant Project Manager and organize into specialized groups to study (in this case)
aerodynamics, attitude control, communications, human factors and science, mission design,
power, propulsion, and structures and thermal control.
At the end of the semester the students deliver a formal presentation of their results. Besides
this report, the class provides an appendix, which provides detailed analyses of their methods and
trades studies.
The quality of the work in this report is consistent with the high standards of the aerospace
industry. The students who participated in this study have demonstrated that they have mastered
the fundamentals of astronautics, have learned to work efficiently as a team, and have discovered
innovative ways to achieve the goals of this project.
In this particular project, the students were challenged to minimize the cost of a human
mission to Ceres (the largest asteroid in the asteroid belt) subject to the following constraints.
Prior to launching a crew of 6 people (3 men and 3 women), two in-situ propellant production
facilities must land on Ceres and produce, not only propellant for the return trip, but also water
and oxygen for the crew to use after arriving at Ceres. The crew outbound voyage and the
return trip should each take less than 2 years each. During the crew interplanetary transfer,
artificial gravity at 0.38 g (equivalent to the gravity on Mars) should be provided. During the
stay time (of 3 months to 2 years) the crew explores Ceres in pressurized rovers that are capable
Project Overview Foreword Page 3
Author: Prof. James Longuski
of ranging over a distance of 765 km in one week. The crew should be safely returned to Earth
(in good health) with a probability of 95%.
I believe this design team rose to the occasion to produce an important feasibility study. The
leadership of the Project Manager and Assistant Project Manager, as well as the outstanding
cooperation of the team members, were key elements in the success of their project. They have
every right to feel proud of their accomplishment and I am proud of them.
James M. Longuski, Ph.D.
Professor of Aeronautics and Astronautics
Purdue University
April 1, 2011
Project Overview Acknowledgments Page 4
Author: Courtney McManus
1.2. Acknowledgements
We thank our professor, Dr. Longuski, and our TA, Frank Laipert, for their guidance and
support throughout the entirety of this project. We would also like to thank the many faculty
members, graduate students, and external contacts who have helped us tremendously with this
design.
Outside Resources
Dr. Boris Yendler – Assistance with structural analysis and thermal control
Dr. David Minton – Assistance with Ceres scientific concerns
Dr. Cary Mitchell – Assistance with hydroponics and human factors issues
Dr. John Rusek – Assistance with engine and propellant concerns
Dr. Charles Koursgrill – Assistance with development of rover suspension systems
AAE Faculty Assistance
Professor Anderson – Assistance with propulsion concerns
Professor Filmer – Assistance with communications concern
Professor Heister – Assistance with propulsion concerns
Professor Howell – Assistance with satellite trajectory analysis
Professor Marais – Assistance with risk analysis
Graduate Student Assistance
Michael Mueterthies – Assistance with trajectory design
Christopher Spreen – Assistance with trajectory design
Project Overview Project Team Page 5
Author: Courtney McManus
1.3. Project Team
Table 1.3-1: The Project Team
Team Member Group
Justin Axsom Communications, Group Lead
Drew Crenwelge Power
Andrew Curtiss Structures and Thermal, Group Lead
Sarah Jo DeFini Communications
Jared Dietrich Propulsion
Anthony D’Mello Communications
Frank (Trey) Fortunato Attitude and Control, Group Lead
Paul Frakes Attitude and Control
Austin Hasse Aerodynamics
Evan Helmeid Assistant Project Manager, Mission Design
Michael Hill Propulsion, Group Lead
Matthew Hill Power
Leonard Jackson Structures and Thermal
Graham Johnson Mission Design, Group Lead
Alex Kreul Structures and Thermal
Joel Lau Power
Chris Luken Attitude and Control
Kimberly Madden Structures and Thermal
Courtney McManus Project Manager
Alex Park Power, Group Lead
Devon Parkos Aerodynamics
Zachary Richardson Human Factors and Science, Group Lead
Jillian Roberts Human Factors and Science
Alexander Roth Aerodynamics, Group Lead
Megan Sanders Mission Design
David Schafer Attitude and Control
Trieste Signorino Mission Design
Elle Stephan Power
Benjamin Stirgwolt Human Factors and Science
Kyle Svejcar Propulsion
Sonia Teran Mission Design
Brendon White Human Factors and Science
David Wyant Propulsion
Project Overview Vision Statement Page 6
Author: Graham Johnson
1.4. Vision Statement
Behind the Vision
Having a vision is something that allows one to see a world without their own eyes and to
understand how they can interact with it. To be able to design such an intricate and creative
project, we have all had to have some personal vision of what we see and perceive the solution of
our mission to be. By displaying and teaching each of our personal visions to one another
throughout the design process, we constructed an even larger collective vision from all of our
ideas becoming one whole solution.
Every time we sit back and look at all the concepts we have learned, ideas we have shared,
and dead ends we have hit, we think of how incredible it is that we could all come together to
build such a remarkable project. One cannot explain in words how exhilarating, enlightening,
and humbling it has been to be able to create this amazing project from the combined vision of
everyone involved.
Project Introduction Page 7
2. Project Introduction
Project Introduction Page 8
2.1. Report Organization
This report was compiled after a nine-week design process completed by the students of Team
Vision. This report contains information on our final design, as well as the road we took to get
there. The main body of this report contains a full explanation of the final design of our project,
such as a cost and risk analysis, the technical systems of the vehicles, the trajectory design, and
many other high-level details of the final design. The appendices of this report contain all of the
analytical work we completed to achieve our final design, as well as a record of designs that
analyzed and eventually discarded. Please refer to the Table of Contents at the beginning of this
report to help navigate through its contents.
In order to accomplish much of the analysis, our team wrote a plethora of computer codes. An
electronic copy of each of the codes is found on the team’s website and in the accompanying CD,
while a user’s guide to each code is located in the Appendix of this report. Our team website is
located at the following link:
https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2011/spring.
Project Introduction Page 9
2.2. Project Objective
The objective of Project Vision is to demonstrate the feasibility of a crewed mission to the
dwarf planet Ceres, while minimizing the overall cost of the mission and assuring a safe return of
the crew. This objective must be met while also meeting all of the Mission Design Requirements
set forth at the beginning of the semester by Professor Longuski. More information on these
Mission Design Requirements is outlined in the following section of this report.
This project serves as a capstone to the education the team has received here at Purdue
University. As such, there was a focus throughout the semester all aspects of a well rounded
engineer. Aside from the obvious focus on technical knowledge and analysis, the team also
focused on presentation skills, written verbal skills, interpersonal relationships, and working
through a multidisciplinary design processes.
This report is the proud result of a semester’s worth of work put forth by 33 senior students in
Aeronautical and Astronautical Engineering at Purdue University in the spring of 2011.
Throughout this semester, we have spent countless hours working hard to design Project Vision
to the best of our abilities. We hope that this hard work and dedication shines through in the
pages of this report.
Project Introduction Page 10
2.3. Mission Design Requirements
Our final design meets the many Mission Design Requirements (MDR) that were set forth at
the beginning of the project. We break these requirements down into the following categories:
logistics, launch, interplanetary transfers, landing, surface operations, in-situ propellant
production (ISPP), exploration and science, and return to Earth. These categories and the
reasoning behind the requirements are explained in the following sections. For a complete
breakdown of the MDRs, please see section A.2.3 of the appendix of this report.
2.3.1 Logistics Requirements
The crew for our mission consists of 6 middle aged people (3 women and 3 men). Every
technology used is space-rated before the first human flight launches. (Please see section 4.1.1
for more information on technology readiness.) We maintain continuous 2-way high definition
video communication between the crew and Earth at all times during the mission.
2.3.2 Launch Requirements
We launch the first cargo flight no earlier than the year 2020. Before the crew can leave
Earth, we ensure that the ISPP stations are fully fueled and completely redundant. The
interplanetary transfer to Ceres lasts less than two years.
2.3.3 Landing on Ceres Requirements
All vehicles landing on Ceres have the capability of hovering over the surface for 60 seconds
before finally landing. This hovering time helps to ensure that the vehicle lands in a safe area and
in a stable configuration.
Project Introduction Page 11
2.3.4 Surface Operations Requirements
The crew stays on Ceres for a minimum of three months and a maximum of two years. This
limit ensures sufficient time for exploration, while not exposing the crew to excessive amounts
of radiation and other hazards of spaceflight. The crew lands at ISPP Station 1 and, midway
through the stay, transfers to the second station. We perform this transfer to allow the crew to
explore as much of the surface of Ceres as possible.
2.3.5 In-Situ Propellant Production Requirements
The vehicle used for crew transfer to Ceres carries only enough propellant for the outbound
journey. We place two ISPP facilities on the surface of Ceres, located at antipodes of each other.
The ISPP stations collect the rocks from the surface, heat the rocks and use electrolysis to release
and collect the resulting propellant. Please see section 4.1.2 of this report for more information
on the Rock Model used for this project.
In addition to creating the liquid hydrogen and liquid oxygen used for propulsion, the stations
create the necessary amounts of breathe-able oxygen and water for the crew to use during the
stay and return journey. We launch the ISPP facilities to Ceres before the first human flight and
ensure the facilities are finished with production by the time the crew leaves.
2.3.6 Exploration and Scientific Requirements
We accomplish scientific exploration using two Exploration Rovers which are stronger, faster,
and safer than crewmembers in spacesuits. These rovers are pressurized to allow the crew to
work in a shirtsleeve environment during the exploration sorties. The rovers are equipped with
sufficient scientific tools, as well as actuators which are dexterous enough to pick up a dime yet
Project Introduction Page 12
strong enough to lift a boulder. The rovers are capable of traveling with two crewmembers one-
quarter the circumference of the planet and returning one ton of rock samples to the crew habitat
every week, as measured in Earth time.
A rescue rover is available on the surface should the need arise to quickly extract an
Exploration Rover crew from a vehicle in distress. The Rescue Rover carries a crew of two, with
the possibility of transporting four additional astronauts. Spacesuits are used only as a safety
precaution. For more information on the use of spacesuits, please see section 4.1.3 of this report.
While the crew is on the surface, they place four seismometers as widely spaced as possible.
A test mass is crashed into the planet from orbit to calibrate the seismic stations and to obtain
information about the core structure of Ceres.
2.3.7 Earth Return Requirements
The crew returns to Earth less than two years after leaving Ceres and brings with them a 1 ton
sample of rocks collected from the surface of Ceres. The crew returns in good health with a
probability greater than 95%.
Project Introduction Page 13
2.4. Design Process
Over the course of this feasibility study, we analyzed and weighed many options and
iterations of design concepts before the final design was selected. We explain those designs and
options which were not incorporated into the final design in greater detail in the appendix
accompanying this report. We outline the design process used during the semester is outlined in
the following paragraphs.
On the first day of class, the team nominated and elected a Project Manager (PM) and
Assistant Project Manager (APM). To help facilitate the organization of our team, each team
member is placed into a specific technical group. These groups consist of aerodynamics, attitude
and control, communication, human factors and science, mission design, power, propulsion, and
structures and thermal. Later, we further divided our team into vehicle groups to ensure that all
of technical aspects of each vehicle were covered. Both the technical and vehicle groups had a
Group Leads to facilitate communication within the team.
Soon after being elected, the PM and APM created a schedule for the team to follow
throughout the semester. This schedule included a few milestones:
Preliminary Design Review
Critical Design Review
Incremental Design Freeze
Final Design Review
These milestones are discussed in further detail in the following sections.
Project Introduction Page 14
2.4.1 Preliminary Design Phase
After being given the initial design requirements, we perform much brainstorming and
discussion to try and find many as many solutions as possible to this open-ended design problem.
However, we do not make any decisions about the final design at this point. We do not make
such decisions until we reach the Preliminary Design Review (PDR) which is scheduled for five
weeks after the design study begins. This scheduling is done to provide ample time for the team
to come up with as many feasible solutions as possible. The main data that concerns us are the
mass, power, and volume values of the vehicles and the vehicle subsystems.
The team addresses the following topics at the PDR:
1) Basic Crew Timeline – How long will it take to transfer to and from Ceres? What is our
maximum stay time? In-Situ Propellant Production (ISPP) stations must know how much
propellant to create.
2) Crew Capsule – Does the capsule remain attached to the Crew Transfer Vehicle (CTV)
during the entire journey or rendezvous in low Earth orbit?
3) Communication Satellites – How many satellites are needed? Will the satellites use
optical or RF communication links?
4) Rovers – How will the rovers move? Will they hop, hover, wheel? How will the rovers
travel from ISPP Station 1 to ISPP Station 2?
5) Crew Transfer Vehicle – What is the basic configuration of the Crew Transfer Vehicle?
How many engines are needed and where? How is it constructed? How will it transfer
from ISPP Station 1 to ISPP Station 2?
6) Supply Transfer Vehicle – What is the basic configuration of the Supply Transfer
Vehicle? How will the cargo land on the surface? How long will it take to transfer?
Project Introduction Page 15
7) ISPP Stations – How will the stations collect the rocks? How will it create the propellant
and consumables? How much of each product must be made?
8) Engines – What types of engines will be needed on each vehicle? How many?
The Preliminary Design effort of our project takes us through the first five weeks of the
semester and is followed by the Critical Design phase.
Project Introduction Page 16
2.4.2 Critical Design Phase
After making initial decisions in the Preliminary Design Review, we enter the Critical design
phase. During this phase, we complete more analysis on our initial design and begin to take the
designs into the subsystem level. We solidify the basic form and functions of our vehicles, as
well as other critical factors.
We conduct a Critical Design Review (CDR) at the culmination of this phase. The CDR
consists of presentations given by the Vehicle Leads about the overall concepts of the vehicles,
including mass, power and volume numbers, as well as the means of propulsion, power
generation, communication systems, etc. Throughout the presentations, the team attempts to find
holes in the designs and other ways in which the design can be improved. At the end of the
presentations, each Vehicle Lead explains to the team what the vehicle group’s forward work
will be and how the changes of the vehicle parameters may affect other vehicles. Some other
topics of consideration are:
1. Communication Satellites – Is the relay satellite needed? What sort of thermal control will
be used on the satellites? What is the exact configuration of the solar panels and the
system bus?
2. ISPP Facilities – How will the tanks connect to the oven and electrolyzer? How will we
organize the storage tanks on the ground? What is the most efficient size of the
Harvesters? How will the facilities be deployed on the surface?
3. Crew Capsule – Where is the hatch located on the capsule? How big must the ballute
tethers be? How will the capsule rendezvous with the Crew Transfer Vehicle? Where are
the attitude and control thrusters located? Where are the sample return rocks housed?
Project Introduction Page 17
4. Rovers – Will the exploration rovers use treads or wheels? Where will the wheels be
placed? How are the rovers deployed on the surface after landing? Will the rovers be
attached to the CTV during the transfer to ISPP Station 2? How do the rovers attach to the
CTV for crew ingress? What sort of attitude control is on the Rescue Rover to use during
flight?
5. Supply Transfer Vehicle – What is the mass of the propellant needed for the outbound
trajectory? How do the modules land on the surface of Ceres?
6. Supply Launch Vehicles – How many launches will be needed? What is the chronology of
the launches? Is it possible to use some of the empty Ares V upper stage tanks for the
ISPP stations?
7. Crew Launch Vehicles – How will we pack the Crew Transfer Vehicle into the launch
vehicles? How many launches are needed? How long will construction of the CTV take in
low Earth orbit?
8. Crew Transfer Vehicle – Where will the crew capsule be located during all phases of the
flight? How does the CTV connect to the rovers for ingress? How many docking ports
will the CTV have? How will the engines be protected during aerocapture? How will the
CTV be assembled in orbit?
The Critical Design phase lasts us two weeks and is followed by the Final Design phase.
Project Introduction Page 18
2.4.4 Final Design Phase and Design Freeze
The Final Design Phase is where we ensure the subsystems of the vehicles are fully defined
and that the interfaces between these subsystems, as well as between the vehicles themselves, are
feasible and defined. This phase takes us through much analysis as we check and re-check our
work to ensure that the final design is a good one.
Because so many of our systems are found to be interdependent, we implement a staggered
design freeze to finalize the design of our vehicles. These interdependencies can be seen in the
following figure.
Figure 2.4.4-1: Dependencies of vehicles and mission aspects
Once we establish the interdependencies of the vehicles, we are able to establish the order in
which the vehicles are frozen. Starting from the bottom of Fig. 2.4.4-1 and working our way up,
Mission Timeline
Number of launches
required
STV CTV
Tanks Rovers
Halo
Satellites
ISPP
Stations
Propellant
mass
Crew
Capsule
Tanks
Propellant
mass
Relay
Satellite
Project Introduction Page 19
we see that first the masses of the propellant for the interplanetary transfers of the STV and CTV
need to be determined. Once we have these values, we can size the propellant tanks
appropriately. Next, it is necessary to freeze the dimensions of the tanks, as well as the designs of
the ISPP stations, the halo satellites, and the rovers so that we can determine how much mass the
STV must transfer to Ceres, as well as the volume dimensions of all of the cargo. Similarly, we
freeze the values of the crew capsule and the propellant tanks for the CTV to understand the total
mass of the CTV.
The keen observer may, at this point, wonder how it is possible to determine the mass of the
propellant needed for the CTV and STV if the total vehicle masses are not known. To account
for this problem, we use an estimate of the masses of the vehicles, then add an extra 15% of the
propellant mass to account for any extra mass added to the vehicles as the designs are finalized.
(This extra mass also accounts for estimating burn arcs for the engines during STV and CTV
transfers. See sections 5.2 and 5.4 for more information on the CTV and STV trajectory
assumptions).
Once we freeze the parameters and dimensions of the STV and CTV, we determine how
many launches are required to bring all of the vehicle components to low Earth orbit (LEO). We
assume an estimate of one launch per every two months during construction and are then able to
determine how long construction in LEO will take. With this duration and knowing how long the
interplanetary transfers take, we make the timeline for our overall mission. From this timeline,
we know whether or not we need the relay satellite based on planetary alignment over the
duration of our mission.
By implementing this staggered design freeze, we are able minimize the impact that
significant changes in one vehicle have on other vehicles. All vehicles are frozen with a specified
Project Introduction Page 20
“fudge factor” of between 5-15% of the overall vehicle mass to account for any errors made or
systems not accounted for.
The final design freeze takes us through two weeks of the project, with the staggered design
freeze lasting one week. These two weeks culminate in the Final Design Review (FDR), which
we hold to ensure that nothing has been overlooked in the design process and that the overall
architecture of our mission is both feasible and accurate..
After the FDR, the design portion of our project is complete, no more analysis is done, and
the final parameters and dimensions of the design are frozen (that is, they can no longer be
changed). It is this final design which is presented in this report. For more information on other
designs which were considered and eventually discarded throughout the design process, please
see the appendix of this report.
Mission Overview and Timeline Page 21
3. Mission Overview and Timeline
Mission Overview and Timeline Quick Reference Guide for Vehicle Specifications Page 22
Author: Evan Helmeid
Co-Author: Frank Fortunato
3.1. Quick Reference for Vehicle Specifications
In this section, we present a general breakdown of the different vehicles and relevant
specifications to serve as a convenient reference guide. These specifications include overall
mass, power, and volume of each vehicle, as well as applicable dimensions and rates.
Supply Launch Vehicle
Table 3.1-1 Launch vehicle type and quantity needed for SLV
Launch No. Component Cargo
STV 1 1 Outer module Reactor
2 Center module ISPP, harvesters, rovers, food
3-5 Outer module LOX, low thrust engine
6-7 Outer module LH2
8-9 Outer jettisoned module LOX
STV 2 1 Outer module Reactor
2 Center module ISPP, harvesters, food
3-5 Outer module LOX, low thrust engine
6-7 Outer module LH2
8-9 Outer jettisoned module LOX
10 Outside components Halo satellites, telemetry dish
Total 19 - -
Supply Transfer Vehicle
Table 3.1-2 Specifications summary of the supply transfer vehicles
Wet mass (T) Dry mass (T) Power (kW) Volume (m3)
STV1 1068 127.3 1236 7818
STV2 1033 131.8 1236 7818
Mission Overview and Timeline Quick Reference Guide for Vehicle Specifications Page 23
Author: Evan Helmeid
Co-Author: Frank Fortunato
Crew Launch Vehicle
Table 3.1-3 Summary of the launch vehicle type and quantity needed for CLV
Launch No. Vehicle Component Cargo IMLEO (T)
CTV 1 Ares V Dry masses Crew quarters 129.5
2-4 Ares V Primary 1-3 Primary tanks and engines 390.3
5-7 Ares V Earth depart. 1-3 Earth depart. tanks and engines 561.7
Crew 1 Ares I Crew vehicle Crew capsule, crew 9.834
Total 8 - - - 1091
Crew Transfer Vehicle
Table 3.1-4 Summary of the CTV specifications
Specification Value Units
Max wet mass 1112 T
Max dry mass 168.8 T
Min dry mass 160.6 T
Post-aerocapture mass 116.9 T
Power usage with low thrust 2020 kW
Power usage at Ceres 59.76 kW
Vehicle volume (internal) 391.3 m3
Ave spin rate 2 rpm
Ave simulated gravity 3.711 m/s2
Crew Capsule
Table 3.1-5 Summary of the Crew Capsule specifications
Specification Value Units
Wet mass 9.834 T
Dry mass 8.932 T
Earth landing mass 8.118 T
Power required 1.889 kW
Pressurized internal volume 27.46 m3
Inflated ballute volume 9.809e4 m3
Mission Overview and Timeline Quick Reference Guide for Vehicle Specifications Page 24
Author: Evan Helmeid
Co-Author: Frank Fortunato
ISPP Stations
Table 3.1-6 Nominal technical specifications for one of the identical ISPP stations
Specification Value Units
System mass (empty) 27.74 T
*Water produced 46.39 T
*Oxygen (g) produced 6.825 T
*Oxygen (l) produced 538.8 T
*Hydrogen produced 118.5 T
Power usage 582.9 kW
Total tank volume 2330 m3
Total system volume 2500 m3
Regolith collection rate 62.00 T/day
Total regolith collected 5.108e4 T
2.421e4 m3
Production time 2.256 years
* Minimum amount required to complete the mission
Exploration and Rescue Rovers
Table 3.1-7 Specifications for one of the twin exploration rovers and the rescue rover
Specification
Exploration
Rover
Rescue
Rover Units
Wet mass 13.29 9.428 T
Dry mass 11.50 6.413 T
Power requirement 25.22 9.15 kW
Internal living volume 12.00 6.000 m3
External volume 65.74 42.18 m3
Range (round-trip) 1531 1531 km
Top rate of travel 4.000 351.2 m/s
Trip frequency 1 1 trip per week
Life support capabilities 4 6 persons
7 1 day(s)
Mission Overview and Timeline Quick Reference Guide for Vehicle Specifications Page 25
Author: Evan Helmeid
Co-Author: Frank Fortunato
Communication Satellites
Table 3.1-8 Comparison between the halo and the Earth-trailing communication satellites
Specification
Halo
satellite 1
Halo
satellite 2
Earth-trailing
satellite Units
Wet mass 17.47 17.47 6.262 T
Dry mass 12.91 12.91 3.825 T
Power requirement 58.22 58.22 52.00 kW
Volume (packed) 324.8 324.8 461.1 m3
Mean distance from Sun 2.764 2.767 1.000 AU
Delta-V to initialize orbit 0.374 0.251 3.220 km/s
Attitude/control propellant 1.347 1.347 45.00 T (over 5 yrs)
Mission Overview and Timeline Acronyms and Definitions Page 26
Author: Evan Helmeid
3.1.1 Acronyms and Definitions
In an effort to reduce wordiness of sentences, many acronyms and abbreviations are used
throughout this report. Below, we present a list of the most common acronyms and for what they
stand. Some of these are repeated in-text and in the Vehicle Names, section 3.1.2. Please note
that this list is by no means exhaustive, but it provides a solid foundation and serves as a
reference guide.
ALARA as low as reasonably achievable
BFO blood-forming organs
CFR Code of Federal Regulations
CFRP carbon fiber reinforced plastic
CPD crew passive dosimeter
CLV Crew Launch Vehicle
CMG control moment gyroscope
CTV Crew Transfer Vehicle
EMU extravehicular mobility unit
EOS Earth Orbiting Satellite
ETRS Earth Trailing Relay Satellite
EVA extra-vehicular activity
FORSE final orbit raise and stabilization engine
g0 ≡ 9.80665 m/s2, Earth reference acceleration due to gravity
GCR galactic cosmic radiation
HDTV high-definition television
HOS Halo Orbiting Satellite
Mission Overview and Timeline Acronyms and Definitions Page 27
Author: Evan Helmeid
HS heat shield
HSHX heat supply and heat exchanger
Isp specific impulse, where thrust = Isp*g0, [s]
Isp,v specific impulse in a vacuum, where thrustvac = Isp,v*g0, [s]
IMLEO initial mass to low Earth orbit
ISPP In-Situ Propellant Production [Facility]
IVLEO initial volume to low Earth orbit
LCO low-Ceres orbit (50 km)
LED light-emitting diode
LEO low-Earth orbit (350 km)
LH2 liquid hydrogen
LOX liquid oxygen
mpropellant mass of propellant
mwet total wet mass of vehicle, including propellant
MF multi-filtration
MLI multi-layer insulation
MMH monomethyl hydrazine
MPD magnetoplasmadynamic
MRU motion reference unit
NCRP National Council on Radiation Protection Measurements
NTO nitrogen tetroxide
OSHA Occupational Safety and Health Regulations
PBA Portable Breathing Apparatus
Mission Overview and Timeline Acronyms and Definitions Page 28
Author: Evan Helmeid
PFE Portable Fire Extinguisher
PMAD power management and distribution
PRD passive radiation dosimeter
RDA radiation dosimeter assembly
RF radio frequency
SAA South Atlantic Anomaly
SLV Supply Launch Vehicle
SMAD Space Mission Analysis and Design
SPE solar particle event
SSLM solid state light module
STV Supply Transfer Vehicle
Sv Sievert (SI unit of radiation)
T metric ton, unit of measurement (1 Megagram)
T thrust, variable or parameter
T:W thrust:weight ratio (weight determined for respective planetary body)
TRL technology readiness level
TT&C telemetry tracking and control
ULCO ultra-low Ceres orbit (25 km)
V∞ excess velocity of the spacecraft
VCD vapor compression distillation
ΔV change in velocity (scalar)
4G fourth generation
Mission Overview and Timeline Vehicle Names Page 29
Author: Evan Helmeid
3.1.2. Vehicle Names
We name all major vehicles in the project based upon mythology and acronyms. The
following list contains the names we chose, as well as a brief description or reason for the
choice.
Supply Transfer Vehicles
Cassiopeia and Cepheus. In Greek mythology, Cepheus was married to Cassiopeia, the
daughter of Adromeda. These two names are also constellations.
Crew Transfer Vehicle
Damocles. In Greek mythology, Dionysius offered his throne to Damocles, a courtier in the
court of Dionysius II, Tyrant of Syracuse. Damocles readily accepted, only to realize that
hanging over his newly-acquired throne was a sword, suspended from the pommel by a single
horse’s hair which could break at any moment and kill him. Suddenly realizing his precarious
situation, Damocles begged Dionysius to take back the throne. We feel the name is appropriate
considering that the CTV employs a tether design; if the tether were to break, the crew would be
lost forever.
ISPP Stations
APES 1 and APES 2. APES stands for Automated Propellant Extraction Station, referring to
the ability of the ISPP stations to extract propellant from the surface material of Ceres.
Exploration Rovers
Castor and Pollux. In Greek and Roman mythology, these twins were patrons of sailors, just
as our rovers protect our astronauts, who are our sailors of the stars. These are also names of
stars.
Mission Overview and Timeline Vehicle Names Page 30
Author: Evan Helmeid
Rescue Rover
SPRINT. SPRINT stands for Speedy Protector Rescues IN Time, referring to the quickness
with which the rescue rover saves the crew in case of a disaster.
Communication Satellites
ECCO 1, ECCO 2, and ECCO Base. ECCO is an acronym for Earth-Ceres Communication
Orbiters, referring to the satellites serving as a communications link between the dwarf planet
and the home planet. We refer to the halo satellites in L1 and L2 orbits as ECCO 1 and ECCO 2,
respectively. ECCO Base is the Earth-trailing relay satellite.
Crew Capsule
ARC. ARC stands for Atmospheric Reentry Capsule, referring to the capsule’s ability to
return the crew safely through the high heating experienced upon reentry into Earth’s
atmosphere.
Mission Overview and Timeline Vehicle Overviews Page 31
Author: Megan Sanders
Co-Author: Paul Frakes
3.2 Vehicle Overviews
3.2.1 Supply Launch Vehicle
The function of the Supply Launch Vehicle is to ensure that all of the components necessary
for the Supply Transfer Vehicle are brought to low Earth orbit (LEO). We select the Ares V
launch vehicle to perform the launches due to its large payload capabilities, both in mass and
volume. A total of 19 launches are necessary to bring all of the structural and cargo components
to low Earth orbit.
Figure 3.2.1 – 1 Structural shell, shown in black, contains payload with engine attached
while inside launch vehicle payload shroud, shown in grey.
We place the payload in a structural shell inside of the Ares V payload shroud to allow the
shroud to deploy in its normal manner while preventing the cargo from being loose in space. All
necessary outside components such as engines and connectors are attached to the structural shell
prior to launch. Having as many components as possible already built in or attached will reduce
the chance of failure during the low Earth orbit construction.
By Megan Sanders
Components by Trieste Signorino
Mission Overview and Timeline Vehicle Overviews Page 32
Author: Sonia Teran
3.2.2 Supply Transfer Vehicle
There are two Supply Transfer Vehicles (STVs), STV1 and STV2, that carry slightly different
cargo. The STVs ferry the necessary cargo for Ceres activities, Ceres communication, resupply
consumables, as well as the In-Situ Propellant Production (ISPP) stations. One mission
requirement specifies that the crew is to land at one location on Ceres and then relocate to
another on the opposite side of the planet. At both of these locations, we place an ISPP station
and resupply consumables, such as food. Therefore, both STVs carry one ISPP station and the
same amount of resupply consumables. We carry all of the rovers on STV1. The rovers wait at
ISPP Station 1 until the crew arrives. All rovers transfer autonomously when the crew relocates
to ISPP Station 2. Two Ceres-orbiting communication satellites provide communication between
Ceres and Earth. We transfer both satellites to Ceres on STV2 and later place them in halo orbits
about Ceres for constant contact with Earth.
Due to decisions discussed in Section 3.2.1, STV components are launched using the Ares V
launch vehicle. We design the dimensions of the propellant tanks and the structural shell to use
as much as the extend shroud volume as possible. We ferry all cargo, except the halo satellites,
inside a structural shell. The halo satellites need to be free to place them on the outside of STV2.
Thus, we place the halo satellites on the outside of STV2 because they do not need to land on the
surface of Ceres. Once STV2 arrives in a captured low Ceres orbit (LCO) of 50 km, the halo
satellites detach and go to their respective orbits.
A more detailed configuration of the STV is discussed in Section 5.2.
Mission Overview and Timeline Vehicle Overviews Page 33
Author: Sonia Teran
Supply Transfer Vechi Timeline
STV1 and STV2 have the same general mission timeline. We construct both STVs in a low
Earth orbit of 350 km. Once complete, the STVs perform a ΔV maneuver and place themselves
in their interplanetary trajectories. We perform the ΔV maneuver with chemical engines that use
liquid hydrogen and liquid oxygen propellant (LH2/LOX). The low-thrust
magnetoplasmadynamic (MPD) engines provide constant thrusting after the ΔV maneuver is
complete. At Ceres, we capture the STVs into LCO. At this point we separate the halo satellites
and send them to their orbits. STV1 then lands at the crew’s initial landing location and STV2
lands at the location for ISPP station 2.
Mission Overview and Timeline Vehicle Overviews Page 34
Author: Trieste Signorino
3.2.3 Crew Launch Vehicle
Placing all components of the Crew Transfer Vehicle (CTV) into low Earth orbit (LEO)
requires two different launch vehicles. Our project requirements do not request a custom design
for a launch vehicle; instead we choose to examine commercially available launch vehicles.
Since the crew of six humans will be traveling to Ceres, we have large quantities of both
structural mass and propellant mass that need to be injected into LEO. This injected mass to LEO
(IMLEO) is a key component in the decision of what type of launch vehicle to use. Another
interesting aspect of launching the CTV is the injected volume to LEO (IVLEO). Our design for
the CTV, and its trip to Ceres, requires large propellant tanks for the trip. These propellant tanks
are the main source of both IMLEO and IVLEO for the CTV and are the basis for the launch
vehicle decision. For the CTV structural components as well as the propellant tanks, we select
the Ares V cargo launch vehicle. We must use the extended shroud design for our launches in
order to avoid more launches than necessary.
The extended shroud consists of a payload fairing with a 10m diameter [1]. Of this 10m, 8.8m
are considered usable for our mission. The height of the extended shroud gives nearly 9 extra
meters of room, compared to the baseline shroud. As mentioned previously, the main advantage
of this launch vehicle and extended shroud is its capability to bring 188 metric tons and 1410 m3
of payload into LEO [2]. The baseline shroud can only carry 860 m3 and 143 metric tons [1].
Using the Ares V extended shroud allows us to send up the CTV propellant tanks, structural
components, and various payloads in a total of 7 launches. A detailed discussion on the masses
and volumes for each launch is located in section 5.3.2.
Along with the placing the CTV in LEO, an additional launch is necessary to deliver our 6
member crew to the Crew Transfer Vehicle. Based on the capsule design and the requirement of
Mission Overview and Timeline Vehicle Overviews Page 35
Author: Trieste Signorino
a human-rated launch vehicle, we make the decision to use the Ares 1 crew launch vehicle. The
Ares I is a two-stage rocket from NASA’s Constellation Program, and is designed to launch the
Orion capsule into LEO [3]. This launch vehicle is capable of injecting approximately 52 metric
tons into low Earth orbit. The dimensions of our Crew Capsule, as well as the mass and volume,
are all comparable to the Orion capsule and therefore fit within the launch capabilities of an Ares
I. A more detailed description of the mass and volume of the Crew Capsule can be found in
section 3.2.5.
Mission Overview and Timeline Vehicle Overviews Page 36
Author: Trieste Signorino
References
[1] Creech, S., “Ares V: New Opportunties for Scientific Payloads,” NASA APO-1052, 2009.
[2] “Overview: Ares V Cargo Launch Vehicle,” Constellation Program ,
URL: http://www.nasa.gov/mission_pages/constellation/ares/aresV/index.html [cited 24
March 2011].
[3] “Constellation Program: America’s Fleet of Next-Generation Launch Vehicles The Ares I
Crew Launch Vehicle,” NASA Facts,
URL: http://www.nasa.gov/pdf/366590main_Ares_I_FS.pdf [cited 24 March 2011].
Mission Overview and Timeline Vehicle Overviews Page 37
Author: Trey Fortunato, Christopher Luken
3.2.4. Crew Transfer Vehicle
Our primary purpose for this mission is to send a six member crew to the dwarf planet Ceres.
The Crew Transfer Vehicle (CTV) accomplishes the daunting task of interplanetary flight. The
mission requirements include a slew of other tasks that we discuss in this section. In short, the
CTV makes the journey to the asteroid belt, lands on Ceres, and returns to Earth where we store
it for future use.
3.2.4.1. Crew Transfer Vehicle Overview
Several systems comprise and configure the Crew Transfer Vehicle, which carry out each
mission requirement. We design the CTV to accomplish seven incredible feats throughout its
mission listed below:
1) Human Interplanetary Travel
2) Simulated Gravity
3) Transmit and Receive HDTV
4) Land and Refuel on an Asteroid
5) Carry Science Material
6) Return to Earth
7) Serve for Possible Future Missions
List items one through three require specific components and implementation methods on our
vehicle. Given the liquid chemical propulsion systems on board, we make round trip
interplanetary travel possible. The vehicle uses a constant low thrust heliocentric transfer orbit
during its journey to Ceres in between planetary departure and arrival burns. Chemical engines
provide the vehicle with acceleration needed for the planetary escape and transfer burns. An
array of electric motors provides constant low thrust during the transfer orbit. The crew habitat
extends outward on a tether system while the entire vehicle is spinning to create an internal
Mission Overview and Timeline Vehicle Overviews Page 38
Author: Trey Fortunato, Christopher Luken
centripetal force for artificial gravity. We maintain continuous High Definition Television
(HDTV) communication using both radio and optical signals throughout the entirety of the
mission while the crew is on board the CTV.
The remaining feats listed above require a compatible configuration consisting of various
systems. Once the Crew Transfer Vehicle arrives at Ceres, it lands, hovers, and refills the
propellant tanks. Rovers collect science material and load it into the capsule storage bay while
the crew is on the surface of the asteroid. The crew transfer vehicle then transfers to a second
propellant production station where the crew collects and stores more science material. Once
Ceres operations are complete, we fully refuel the CTV primary tanks and collect other crew
related consumables for the return journey to Earth. Prior to our aerocapture maneuver, the crew
moves into the re-entry capsule and separates from the CTV. This allows the CTV and re-entry
capsule to perform aerocapture maneuvers independently. The CTV then undergoes an
aerobraking maneuver that gradually lowers the apogee radius. The FORSE fires to circularize
the orbit, inserting the CTV into LEO. The crew transfer vehicle then waits in a parking orbit
around Earth for propellant tanks, crew consumables, and a new power source for its next
interplanetary mission. Figure 3.2.4.1-1 below shows the completed CTV in LEO with the
stowed capsule.
Figure 3.2.4.1-1 Completed Crew Transfer Vehicle with stowed capsule.
By: Alex Roth
Mission Overview and Timeline Vehicle Overviews Page 39
Author: Trey Fortunato, Christopher Luken
3.2.4.2. Crew Transfer Vehicle Mission Timeline
From beginning to end, the Crew Transfer Vehicle fills a variety of roles, which requires a
unique design and implementation method. We break the CTV timeline into four primary
segments, construction, outbound transfer, Ceres surface, and return trip operations. The CTV is
too large to launch directly from Earth’s surface as one vehicle, thus we assemble each piece in
LEO. Upon completion of the construction phase, the CTV fulfills its primary duty of
interplanetary flight. The mission duration then consists of three phases, crew transfer to Ceres,
Ceres operations, and crew return to Earth. As part of returning to Earth, we retain the central
portion of the vehicle for future use, and insert the chassis into a parking orbit. Figures 3.2.4.2-1
through 3.2.4.2-4 display the full timeline followed by detailed descriptions of each phase. A
sequential launch sequence comprises the progression of construction. We attach each
subsequent section to the existing vehicle upon arrival in the construction orbit.
Figure 3.2.4.2-1Crew Transfer Vehicle construction phase.
Mission Overview and Timeline Vehicle Overviews Page 40
Author: Trey Fortunato, Christopher Luken
Figure 3.2.4.2-2 Crew Transfer Vehicle outbound transfer to Ceres.
The Earth-Ceres transfer begins with stowing the crew capsule on top of the vehicle chassis
after the crew has disembarked. We initialize the trip to Ceres when the CTV undergoes an Earth
departure burn followed by a constant low-thrust heliocentric transfer. Low thrust
Magnetoplasmadynamic electric motors provide thrust during this heliocentric spiral. We
perform another high thrust burn upon the arrival at Ceres. Changes in each step of this phase
require several vehicle re-orientations. The final configuration of this phase positions the capsule
on the side in preparation for Ceres surface operations.
Mission Overview and Timeline Vehicle Overviews Page 41
Author: Trey Fortunato, Christopher Luken
Figure 3.2.4.2-3 Crew Transfer Vehicle surface operations on Ceres surface.
At and on Ceres, we attach the capsule to the habitat docking port for regolith collection
purposes. With the capsule positioned correctly on the side of the CTV, the vehicle performs
descent and hovering maneuvers. The CTV lands near the ISPP stations to replenish resources
and to prepare for station transfer.
Figure 3.2.4.2-4 Crew Transfer Vehicle transfer to Earth.
Mission Overview and Timeline Vehicle Overviews Page 42
Author: Jillian Roberts
3.2.5. Crew Atmospheric Re-entry Capsule
The Atmospheric Re-entry Capsule (ARC) has two main purposes. The first is to transport
the astronauts from low Earth orbit (LEO) to the Crew Transfer Vehicle. Second, following
Ceres to Earth transfer, the capsule safely returns the astronauts and Ceres regolith to the surface
of the Earth. We achieve these objectives through a four-phase timeline. Phase I is launch and
rendezvous with the CTV, Phase II is outbound trip and return, Phase III is Earth approach, and
Phase IV is capsule re-entry and recovery.
3.2.5.1 Phase I: Launch and Dock
In Phase I, an Ares I rocket boosts the capsule with astronauts into Low Earth Orbit (LEO).
The astronauts check all systems and search for anomalies which may adversely affect the
mission. If one or more critical systems are compromised, the astronauts can abort the mission
and return to Earth. If all systems are nominal, Mission Control gives the go-ahead for
rendezvous with the Crew Transfer Vehicle (CTV). The capsule docks to the side of the CTV,
and the astronauts climb into the crew habitat, their home and command center for the next four
years. The crew capsule autonomously detaches and re-docks to the top of the CTV stack. We
show this procedure in Fig. 3.2.5.1-1.
Mission Overview and Timeline Vehicle Overviews Page 43
Author: Jillian Roberts
Figure 3.2.5.1-1This figure shows the docking and re-docking procedures of the capsule
between the top of the CTV stack and the side of the crew habitat.
3.2.5.2 Phase II: Outbound and Return
Phase II begins with the CTV escaping Earth and transferring to Ceres. The capsule
autonomously detaches from the top of the CTV stack and re-docks to the side of the crew
habitat for convenient loading of Ceres regolith. This maneuver is Step 2 in Fig. 3.2.5.1-1. The
CTV lands and the astronauts carry out the Ceres surface operations. After liftoff of the CTV
Mission Overview and Timeline Vehicle Overviews Page 44
Author: Jillian Roberts
from Ceres, the capsule detaches once again and docks to the top of the CTV stack in preparation
for transfer back to Earth (Step 3 in Fig. 3.2.5.1-1).
3.2.5.3 Phase III: Earth Approach
After Earth transfer, the Atmospheric Re-entry Capsule autonomously docks to the side of the
crew habitat (Step 4 in Fig. 3.2.5.1-1). As the CTV approaches Earth, the six suited astronauts
climb into the capsule. The astronauts check communication, life support, instruments, and
recovery systems on the capsule while still docked to the CTV. Assuming all systems are
nominal, the docking hatches seal closed and the capsule separates from the CTV.
3.2.5.4 Phase IV: Capsule Re-entry
After the capsule separates, it drops into an independent trajectory from the CTV, speeding
around the Earth. The ballute cap pops open, deploying the inflated ballute to begin the
aerocapture maneuver and acquire low Earth orbit. Upon completion of the aerocapture
maneuver, the tethers are cut and the ballute drifts away. At this point, the capsule has lost
enough velocity to begin its descent into the atmosphere. A second cap explodes open, releasing
the triple parachute. The capsule travels through the atmosphere and safely splashes into the
Pacific Ocean. The aerial recovery team pulls the astronaut crew from the floating capsule and
completes the mission.
Mission Overview and Timeline Vehicle Overviews Page 45
Author: Zachary Richardson
3.2.6 In-Situ Propellant Production Stations
We use two In-Situ Propellant Production (ISPP) stations to generate and store water, liquid
oxygen, and liquid hydrogen for our mission. The Supply Transfer Vehicles (STVs) place their
corresponding ISPP stations at antipodes on the Cerian equator. The stations are exact duplicates
of each other in order to provide redundancy and a factor of safety for the crew. Upon arrival at
Ceres, each station produces the propellant required for both the Crew Transfer Vehicles (CTV)
transfer between the two ISPP stations and the return trip to Earth. They also provide the water
and oxygen needed by the crew to survive day to day activities while the crew is on Ceres and
the return trip. The stations generate enough propellant to supply both Exploration Rovers and
the Rescue Rover for their respective missions on Ceres. The model below displays our complete
ISPP station:
Figure 3.2.6-1 The complete ISPP station show here in its “deployed” or operations mode.
Figure By: Alex Roth
Mission Overview and Timeline Vehicle Overviews Page 46
Author: Zachary Richardson
The ISPP stations travel to Ceres aboard STV-1 and STV-2. Each individual station includes
three Harvesters (two operating on the surface while the third waits in stand-by mode in case of
emergency). These Harvesters collect the surface regolith around the station. Upon the landing of
the STV, the ISPP stations establish communication with Earth via the Halo Satellites. The
collection bin is deployed as soon as the reactor has begun producing power. The Harvesters
begin collecting regolith once communications have been established between the Harvesters,
the ISPP stations and Earth. The central core of the station contains the nuclear reactor,
electrolyzer, conveyor belts, pumps, and condensers. The input conveyor belt delivers regolith to
the oven once the Harvesters begin depositing material in the collection bin. The oven heats the
rocks until the water inside can be extracted as a vapor. From here an electrolysis process
separates the water into hydrogen and oxygen. Holding tanks store the condensed elements.
Once the desired amount of material has been extracted, the stations go into a standby mode
where only the holding tanks require power. The reactor continues to operate and will assist in
powering the propellant transfer from the ISPP station to the Rovers and the CTV. After the crew
embarks on their homeward journey, the ISPP stations will completely shut down with no further
activity. For a more detailed analysis, please continue to section 5.6.
Detailed Vehicle Descriptions Exploration Rovers Page 47
Author: Ben Stirgwolt
3.2.7. Exploration Rovers
Upon arriving at Ceres, the astronauts conduct scientific experiments and explore the dwarf
planet. The astronauts conduct some of the experiments while on their way to and from Ceres,
but they perform the majority of the experiments on the exploration rovers. The mission
requirements for Project Vision require that two exploration rovers, named Castor and Pollux,
simultaneously embark on sorties, allowing for a rescue if one becomes non-functional. Each
exploration rover holds two astronauts, but can hold up to four people if a problem arises in the
other rover. The remaining two astronauts in the CTV monitor the status of each exploration
rover. If a medical emergency arises or both of the exploration rovers become non-functional,
then the two astronauts in the CTV embark on a rescue mission with the rescue rover.
The exploration rovers are required to gather a ton of regolith and to travel a distance of up to
¼ the circumference of Ceres (about 765 km) every week. To meet these mission requirements,
we opt for a wheeled vehicle capable of travelling at speeds of up to 4 m/s, using an internal
combustion engine as the means of propulsion. Each rover also has two robotic arms, one at the
fore and one at the aft of the rover, which are capable of picking up rocks as small as a dime or
as large as a boulder. There are also two rock storage containers located at the fore and aft of the
rover where the regolith is stored for the duration of the mission. Upon completion of a sortie,
the rovers return to the CTV, and the astronauts use the robotic arm to lift the storage box from
the rover, dump the rocks into a pile for sorting, and then replace the empty storage container in
position on the rover. Figure 3.2.7-1 shows the exterior configuration of the exploration rover
with the rock container and the robotic arm in the front of the rover.
Detailed Vehicle Descriptions Exploration Rovers Page 48
Author: Ben Stirgwolt
Figure 3.2.7-1 The exploration rover serves as a moving laboratory where astronauts
conduct scientific study, using a wide range of tools and equipment in a “shirt-sleeve”
environment.
The exploration rover is essentially a cylindrical pressure vessel with a temperature-controlled
environment. This “shirt-sleeve” environment makes each sortie move comfortable for the
astronauts as well as easier to conduct experiments since they do not have to wear space suits.
The rovers are equipped with a glove box so that the astronauts can perform experiments on the
rocks without bringing them into the earth-like atmosphere inside of the rover. The rovers have
microscopes, telescopes, and a variety of geologists’ tools. The science equipment and the other
electric equipment on the rovers are supplied with electricity from an internal combustion engine
By: Kim Madden
Detailed Vehicle Descriptions Exploration Rovers Page 49
Author: Ben Stirgwolt
in addition to an electrical generator and three sodium-sulfur batteries. In order to transmit data
back to the CTV and maintain contact with the other astronauts, the exploration rovers
communicate with the halo satellites using radio frequency communication.
Upon completion of a sortie, the exploration rovers return to the CTV. After the rocks are
removed from the storage container, the rovers dock with the CTV using a lifting mechanism and
then one of the two docking ports located on either side of the rover. They remain in this
condition until the next sortie. When the CTV transfers from ISPP station 1 to ISPP station 2
halfway through the mission, the rovers autonomously maneuver themselves to ISPP station 2,
where the process of collection rocks and performing experiments continues. When the crew
leaves Ceres at the end of the mission, we abandon the exploration rovers ISPP station 2.
Mission Overview and Timeline Vehicle Overviews Page 50
Author: Kim Madden
3.2.8 Rescue Rover
The purpose of the rescue rover is to travel to a stranded Exploration Rover to save the crew
members. The Rescue Rover has the capability to go a quarter of the circumference of Ceres
(765 km) and contains enough life support and free space for all 6 crew members to fit inside
comfortably.
Since the Rescue Rover needs to rescue stranded crew members, it must travel quickly; we
have chosen to have the Rescue Rover use rockets to launch into an orbit and land near the
stranded rover. This gets us closer to the stranded crew quickly and efficiently.
We divide the Rescue Rover’s timeline into three stages: Transit, Ceres Surface Missions, and
Seismic Testing. The first stage is the transit. This stage is simply getting the Rescue Rover from
Earth to Ceres. The Rescue Rover launches with the STV, along with the two Exploration
Rovers. Upon landing at Ceres, the STV opens up, and the Rescue Rover has enough propellant
to be remotely operated to the ISPP, where it will be filled with enough propellant for a Rescue.
The Rescue Rover is then remotely controlled back to the CTV, where it waits for a rescue
situation. At this point, crew members stock the rover for a rescue trip and ensure that everything
works properly.
The second stage is the Ceres Surface Missions. This stage includes the rescuing of
potentially stranded crew members, as well as a transit from ISPP Station 1 to Station 2. The two
crew members remaining in the CTV enter the Rescue Rover after a distress call is received. The
Rescue Rover drives a few meters away from the CTV, and then launches into a specific
trajectory depending on the distance of the struggling Exploration Rover. The Rescue Rover has
the capability to hover for 60 seconds before landing, so it can maneuver in close to the
Exploration Rover. Upon landing, the Rescue Rover drives up to the stranded rover, where they
Mission Overview and Timeline Vehicle Overviews Page 51
Author: Kim Madden
dock together to save the explorers. Once the crew is safely inside the Rescue Rover, it drives a
few meters away from the now-empty Exploration Rover, and launches again into its specific
trajectory towards the CTV. Upon landing, the Rescue Rover hovers for up to 60 seconds to get
as close the CTV as is safe. The Rover then drives to dock with the CTV, where the crew can
safely egress and restock for the next rescuing mission. This process can be repeated as many
times as necessary (but we hope it never happens!).
The next part of the second stage is the transit from ISPP station 1 to 2. The Rescue Rover
transfers to the other ISPP station by remote control. Once the CTV lands at station 2, it tells the
rovers that it is time for them to also transfer to station 2. The rover does one hop with a longer
orbit time to make it to the other station, where it safely lands as previously discussed. The CTV
remotely controls the Rescue Rover to receive more propellant from ISPP Station 2 and then
attaches itself back to the CTV. The crew restocks the Rescue Rover for future use, and waits for
its shining moment.
The last stage of the Rescue Rover’s life is the Seismic Testing. This occurs after the crew
launches back into a Low Ceres Orbit and it about to head home. The Rescue Rover engine burns
for 21 seconds to a height of 26.5 km, where it turns around and propels itself towards the
surface. This creates a large impact, and seismic monitors pick it up to determine the
composition of Ceres (see Section 5.8.10. for a more in depth description of the seismic testing).
Table 3.2.8-1 shows the final mass, power and volume numbers for the Rescue Rover. These
values include a 7% increase from the actual calculated values as a buffer. Figure 3.2.8-1 shows
a model of the Rescue Rover. The analysis for the Rescue Rover can be found in Section 5.8.
Mission Overview and Timeline Vehicle Overviews Page 52
Author: Kim Madden
Table 3.2.8-1: Rescue Rover Mass, Power and Volume Results
Technical Group Injected Mass
to LEO, kg
Additional
Mass, kg
Power, kW Component
Volume, m3
Injected Volume
to LEO, m3
Human
Factors/Science
503.00 591.00 3.12 13.12 0.10
Structures &
Thermal Control
5,189.80 0 0.02 3.19 39.00
Propulsion 124.95 2,127.71 4.66 0 6.61
Communication 16.29 3.60 0.72 0 0.05
Attitude Control 8.00 95.00 0.03 0.02 0
Power 151.71 0 0 0.27 0
TOTAL 6,413.31 3,014.52 9.15 24.84 42.18
Figure 3.2.8-1: This is a model of the Rescue Rover.
Mission Overview and Timeline Vehicle Overviews Page 53
Author: Tony D’Mello
3.2.9 Communications Network
Our communications network requires that we maintain continuous two-way high definition
television (HDTV) communication between each crew member and Earth during the mission.
We transmit telemetry, tracking, and control data (TT&C) as well, but these data amounts are
quite small compared to an HDTV signal, so for simplicity we reference data rates in terms of
the number of HDTV channels.
3.2.9.1 General Communications Network
3.2.9.1.1 Earth Focused Network
We expect to receive nine high definition television (HDTV) channels from the crew at Ceres,
one from each crew member and one from each rover. We also expect to send six HDTV
channels to Ceres, one for each crew member.
We send and receive information from ground stations on the Earth. These stations currently
exist at universities. Using radio frequencies (RF), they communicate with Earth orbiting
satellites (EOS).
The three EOS are in geostationary orbit, each at equal angular distance from one another.
When one receives a signal from the Earth ground station, it will decode the RF message and
then transcode it into an optical message which will be sent to either the Earth Trailing Relay
Satellite (ETRS) or directly to a Halo Orbiting Satellite (HOS). The reverse is true for when the
EOS receive information from Ceres.
Mission Overview and Timeline Vehicle Overviews Page 54
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Figure 3.2.9.1.1-1 illustrates a data transfer from Earth to Ceres. We send the information, in
the form of an RF signal, from a ground station to the EOS currently in communication with
Ceres. The EOS sends the signal to the HOS or ETRS depending on the conditions.
Figure 3.2.9.1.1-1 We send a message from Earth to Ceres. (Image is not to scale.)
3.2.9.1.2 Intermediate Transfer
Once we place the HOS in their orbits, Earth can communicate directly with Ceres, using an
optical frequency. When Earth enters opposition or conjunction with Ceres, a blackout period
exists. In order to maintain continuous communication, during these moments, the ETRS
communicates between the EOS and the HOS via optical communication. In addition to the
blackout periods, we communicate through the ETRS when it is closer to Ceres than Earth,
thereby limiting the power required by the HOS.
Figure 3.2.9.1.2-1 shows the two possible paths for Ceres/Earth communication. We
communicate via path (1) when Earth is close to Ceres. We communicate via path (2) when the
ETRS is closer to Ceres or during a blackout period.
Mission Overview and Timeline Vehicle Overviews Page 55
Author: Tony D’Mello
Figure 3.2.9.1.2-1 We have two methods for sending information, (1) Ceres to Earth
directly and (2) via the ETRS. Orbits are to scale, but celestial bodies and ETRS are not.
3.2.9.1.3 Ceres Focused Network
Communication on Ceres exists between each vehicle and with the Earth. Therefore, a crew
member in Rover 1 can communicate with a crew member in Rover 2. Each uplink and
downlink signal for each vehicle operates at its own frequency ranging from 7-60GHz.
We will place two satellites, HOS, in halo orbit about Ceres, one at each Lagrange point.
Each of these satellites can send and receive an optical signal to Earth or the ETRS.
Communication with the Crew Transfer Vehicle (CTV) occurs via optical communication.
Each halo satellite has an RF transmitting dish for each rover. We require only one RF receiving
Mission Overview and Timeline Vehicle Overviews Page 56
Author: Tony D’Mello
dish to receive the signals from all three rovers. HOS 1 is capable of sending up to nine HDTV
channels to HOS 2 and vice versa via RF during saturation prevention procedures and in the
event of an optical malfunction. Figure 3.2.9.1.3-1 shows all 3 rover signals being sent to a
single HOS. The optical connection from the CTV to the HOS is also displayed, as well as the
connection between the two HOS.
Figure 3.2.9.1.3-1 We send all the Ceres information to a HOS which will then deliver the
information to Earth. Images are not to scale.
Mission Overview and Timeline Vehicle Overviews Page 57
Author: Tony D’Mello
When the CTV does not reside near the In-Situ Propellant Production (ISPP) station, the ISPP
station can send and receive an RF signal corresponding to TT&C by “borrowing” the dishes on
the halo satellite normally reserved for a rover. This system is shown in Fig. 3.2.9.1.3-1 in the
top right where an ISPP station sends its signal to an HOS which then sends the signal to the
other HOS. When the CTV resides near the ISPP station, the ISPP station sends its information
to the CTV using the same RF transmitting and receiving dishes. In the bottom right of Fig.
3.2.9.1.3-1, a harvester sends information to the ISPP station which then sends the signal to the
CTV.
The CTV has an RF transmitting/receiving pair of dishes for communication with the ISPP
stations. Decoding that information, the CTV adds its own data and sends the sum to the HOS
via optical transfer. The CTV receives data from Earth via the same manner.
Each rover has an RF transmitter/receiver pair of dishes. The exploration rover dishes can
send six (four crew members and two outside camera feeds) HDTV channels and can receive
four HDTV channels (crew members). The rescue rover sends seven (six crew and one outside
camera feed) HDTV channels and can receive six HDTV channels (crew members).
3.2.9.1.4 Crew Communication Interface
We assign each crew member a cell phone-like communication device which they carry at all
times. Present day technology already provides more than sufficient devices in the form of
fourth generation (4G) smartphones and tablet PCs.
For enhanced visual communication, we equip each crew vehicle with organic light-emitting
diode (OLED) televisions measured a little over half a meter in size.
Mission Overview and Timeline Vehicle Overviews Page 58
Author: Sarah Jo Defini
3.2.9.2 Supply Transit Communications Network
The Supply Transfer Vehicles will complete a data dump once every month during their
transit to Ceres. However, we will not have a relay satellite in place during the transfer, so the
Supply Transfer Vehicles are equipped to communicate directly with Earth. During the data
dump the telemetry dish on each vehicle aligns itself so that it can transmit data to a visible a
Tracking and Data Relay Satellite, which relays data to the Deep Space Network on Earth. Earth
can then transmit data to each vehicle.
Mission Overview and Timeline Vehicle Overviews Page 59
Author: Justin Axsom
3.2.9.3 Crew Transit Communications Network
We use a variety of communication links to connect with the crew throughout the lifetime of
the mission. One of the mission requirements is maintaining constant two-way HDTV
communication with each crew member at all times. To accomplish this task, we employ
different methods of communication that heavily depend on the distance of the link. Initially, the
crew launches within the crew capsule on top of the crew launch vehicle. Up until the crew
capsule jettisons from the crew launch vehicle, the crew launch vehicle uses a single high-gain
antenna to transmit the HDTV signals to mission control on Earth. The crew launch vehicle also
transmits multiple channels of tracking, telemetry, and control data through a number of
omnidirectional antennas. Then, once the crew capsule separates from the crew launch vehicle, it
transits to rendezvous with the crew transport vehicle. At this time, we use an ultra-high
frequency, phased-array antenna to transmit the HDTV signals and logistical data to the crew
transport vehicle. The crew transport vehicle then relays that data to the NASA tracking and data
relay satellites which send everything back to mission control. This process continues until the
crew transport vehicle reaches Earth escape. The entire process is summarized in Fig. 3.2.9.3-1.
Mission Overview and Timeline Vehicle Overviews Page 60
Author: Justin Axsom
Figure 3.2.9.3-1 A visual summary of the near-Earth communication links with the Crew.
TDRS refers to the NASA tracking and data relay satellites.
Once the crew transport vehicle escapes from Earth, we start using an optical communication
system. The optical communication system will transmit and receive nine HDTV signals in all.
The crew transport vehicle communicates directly with Earth’s optical receiving satellites or with
the Earth-Trailing Relay Satellite until half-way through the journey from Earth to Ceres. At the
half-way point, we point the optical system at the Ceres halo orbiting satellites. We continue to
communicate with the halo satellites for the duration of the transit to Ceres, while on the surface
Mission Overview and Timeline Vehicle Overviews Page 61
Author: Justin Axsom
of Ceres, and half-way through the return trip. At this point, we again redirect the optical system
to point back at Earth’s optical receiving satellites or with the Earth-Trailing Relay Satellite.
Once the crew transport vehicle is captured by Earth, we can communicate with the NASA
tracking and data relay satellites again until the crew capsule departs. Once the crew capsule
departs, we can relay communication to the crew transport vehicle up until Earth re-entry. The
only significant communication black-out period occurs at this point until the crew capsule
reaches a lower velocity and the plasma surrounding the capsule subsides. Finally, a small
omnidirectional antenna transmits a beacon for recovering the crew. We successfully accomplish
the requirement of constant, two-way HDTV communication with the crew at all times except
for the blackout during Earth re-entry.
Mission Overview and Timeline Scientific Overview Page 62
Author: Ben Stirgwolt
Co-Author: Zachary Richardson
3.3 Scientific Overview
During the course of mission, the astronauts perform many experiments covering a wide
range of fields—geology, physics, astronomy, medicine, and the life sciences. The bulk of the
geology experiments are conducted in the exploration rovers. Little is known about the surface
of Ceres. Even less is known about the interior. Thus a parameter of the scientific mission is to
determine the inner composition of the dwarf planet through the placement of seismometers on
the surface. During the regolith collection missions, we place the seismometers equidistant from
one another so that they are approximately 900 km apart. Figure 3.3-1 shows a possible
configuration of the seismometers on the surface of Ceres. Upon the crew’s departure, the
Rescue Rover launches into Ceres at full speed. The resulting impact allows the seismometers to
generate enough data to be transmitted back to Earth for analysis [1].
Figure 3.3-1 The seismometers are placed at an equal distance from each other to allow for
optimal data collection.
Mission Overview and Timeline Scientific Overview Page 63
Author: Ben Stirgwolt
Co-Author: Zachary Richardson
In addition to the seismic experiment, the rovers deploy several experiments to further study
the composition of the Ceres surface. A meteorite experiment studies the small particles that
strike the surface of Ceres, measuring their velocity and direction at the time of impact. An
electrical properties experiment uses transmitting antennae to determine the electrical properties
of the regolith. In addition to these experiments, the astronauts use a variety of geological tools
including heat flow probes, an electromagnetic sounder, spectrometers, microscopes, and other
tools.
For immediate scientific inquiry of the soil composition, there is a glove box in the
exploration rover so astronauts can examine rock samples in their natural environment, without
having to expose the soil to a foreign environment.
With regards to physics and astronomy experiments, the astronauts will use a traverse
gravimeter that will be deployed at several locations on Ceres to make relative gravity
measurements. They also use a small research telescope and an ultraviolet light telescope study
the evolution of galaxies.
Throughout the mission, the astronauts grow a portion of the food they eat. In addition to the
using the crops as a source of food, they conduct experiments to see how plants behave in a
microgravity environment for an extended period of time. Creating a closed loop life support
system allows for the astronauts to study environmental management and control, agriculture,
food processing, diet planning, and waste processing.
Of course the CTV is stored with medical equipment in case of an emergency in addition
to exercise equipment, but we also stock the vehicle with equipment to perform basic research in
human physiology and neurophysiology. Humans have never experienced microgravity for as
long of a time as the astronauts of Project Vision, affording opportunities to explore uncharted
Mission Overview and Timeline Scientific Overview Page 64
Author: Ben Stirgwolt
Co-Author: Zachary Richardson
territories in human physiology. There are many questions that remain regarding basic
understanding of human adaptability to the artificial gravity environment. As the European
Space Agency mentions, areas of interest include: neurophysiology, cardiovascular physiology,
oxygen metabolism, musculoskeletal system, blood formation, pharmacokinetics, radiation
effects, and immunology [2].
Mission Overview and Timeline Scientific Overview Page 65
Author: Ben Stirgwolt
Co-Author: Zachary Richardson
References
[1] Mizutani, Hitoshi. United States(NASA), Japan (JAXA). National Space Science Data
Center., 2005. Web. 1 Apr 2011.
http://nssdc.gsfc.nasa.gov/nmc/spacecraftDisplay.do?id=LUNAR-A>.
[2] Eckart, P., “Science at a Lunar Base,” The Lunar Base Handbook, McGraw-Hill, 1999, p.
500
Mission Overview and Timeline Page 66
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Project Conclusions Page 67
4. Project Conclusions
Project Conclusions General Project Concerns Page 68
Author: Courtney McManus
4.1. General Project Concerns
4.1.1 Assumptions Made
We design our mission to occur no earlier than the year 2020, and as such, we must design a
mission architecture that will not be outdated by the time the mission is implemented. To
accomplish this, we make various assumptions on the funding and political support the project
will receive, as well as the technology which will be readily available when the mission is
implemented.
4.1.1.1 Political and Funding Assumptions
A year before this project was designed, President Obama gave a speech at the Kennedy
Space Center outlining his goals for America‟s human spaceflight program., a vision which
happens to include a human mission to an asteroid. President Obama is quoted as saying:
“Early in the next decade, a set of crewed flights will test and prove the
systems required for exploration beyond low Earth orbit. And by 2025,
we expect a new spacecraft designed for long journeys to allow us to
begin the first-ever crewed missions beyond the Moon into deep space. So
we’ll start – we’ll start by sending astronauts to an asteroid for the first
time in history.” [1]
Using this speech and the general direction of human spaceflight as a guide, we make the
assumption that our mission will have the necessary political backing, and therefore funding, to
accomplish our design as outlined. We do not take into account the politics surrounding such
things as nuclear reactors and the Planetary Protection Act. It is assumed that a sufficient amount
of international cooperation is used so that this exploration mission can benefit all of humanity.
Project Conclusions General Project Concerns Page 69
Author: Courtney McManus
4.1.1.2 Technology Readiness Assumptions
In order to design a mission which will not be technically irrelevant when implemented, we
made assumptions on the readiness level of various technologies. Technology Readiness Levels
(TRL) were first identified by the US Government to provide a means of assessing the maturity
of technologies which are currently under development. These are numbered TRL 1-9, the levels
increase numerically with maturity. Throughout this report, we may to refer to a technology as
having a Technology Readiness Level of a given number. These levels are defined by NASA as:
TRL 1: Basic principles observed and reported
TRL 2: Technology concept and /or application formulated
TRL 3: Analytical and experimental critical function and/or characteristic proof-of-concept
TRL 4: Component/subsystem validation in laboratory environment
TRL 5: System/subsystem/component validation in relevant environment
TRL 6: System/subsystem model or prototyping demonstration in end-to-end environment
TLR 7: System prototyping demonstration in an operational environment
TRL 8: Actual system completed and mission qualified
TRL 9: Actual system mission proven through successful mission operations [2]
For this design, we assume that technology with a TRL of 3 in 2011 will have an acceptable
amount of funding and research appropriated to be at an operational and space-rated TRL by the
time the mission is flown.
4.1.1.3 Launch Vehicle Assumptions
For the purposes of our mission, we assume that the Ares V and Ares I launch vehicles will be
available for use by the time we fly our mission.
Project Conclusions General Project Concerns Page 70
Author: Courtney McManus
References
[1] Obama, Barrack, “Remarks by the President on Space Exploration in the 21st Century,”
Press conference at the John F. Kennedy Space Center, Merritt Island, FL. April, 2010.
[2] Mankins, John C., “Technology Readiness Levels, a White Paper.” Advanced Concepts
Office, Office of Space Access and Technology, NASA. April 6, 1995
Project Conclusions General Project Concerns Page 71
Author: David Wyant
Co-Author: Zachary Richardson
4.1.2 Ceres Regolith Model
Our regolith model combines a number of factors including the coefficient of friction of the
surface regolith, the frequency of large rocks and boulders on the surface of the planetoid, and
the density and water content of the rocks and the regolith itself.
The coefficient of friction we chose for this project was based on a lunar analogue. This
value was selected to be 0.2 [1]. This is the coefficient of friction for the lunar surface as
determined by observing avalanches around craters on the moon. We chose to use this value as
our coefficient of friction as no other true analogue exists. As both bodies are heavily cratered
by micro-meteors, it seems a reasonable assumption that the surface composition will be very
similar.
The next issue we will consider is that of the frequency of large boulders on the planetoid‟s
surface that could inhibit landing and impede the movement of the different Rovers exploring the
surface of Ceres. Using a previously conducted geological survey of the moon, a table addressing
the frequency of different rock sizes around craters can be created [2]. The results of this analysis
follow.
Table 4.1.2-1 Frequency of Boulders on surface of Ceres
Size 2 meter 4 meter 6 meter
Avg. Rock Frequency (per m^2) 1.234x10-3
1.28x10-4
9.2x10-6
Interpreting the results above, we selected a boulder diameter of 2 meters to be the obstacle
the Rovers would be designed to overcome. This choice is due both to two factors: the higher
frequency of this type of rock and the assumption that boulders of larger diameters will either be
able to be driven around or have an incident slope of 45 degrees or less. The first concern
Project Conclusions General Project Concerns Page 72
Author: David Wyant
Co-Author: Zachary Richardson
represents a higher likelihood of encountering an obstacle that cannot be maneuvered around.
The second issue represents another design criterion for the rovers: the ability to climb a 45
degree incline.
Due to its position in the Asteroid Belt and its apparent lack of geothermal activity, Ceres is
assumed to be relatively unchanged in its elemental composition from its creation. This lack of
geothermal activity has led scientists to believe that the majority of the planet, including the
surface, consists of some of the solar system's earliest known asteroid compositions, CI
chondrite. The same lack of geothermal activity results in no mountain ranges on the planet as
well as no way for the scars left behind by meteor impacts to be covered. The eons of meteoroids
craters create a surface of flat plains with huge crater ridges and drastic drop offs. Our own Dr.
Minton believes that Ceres consists of such a rock type. CI-chondrite has a high water content
which makes the surface an ideal place for an electrolysis based In Situ Propellant Production
(ISPP) facility. The density is one of the lowest among asteroids because of its high water
content. Current estimates place the density of the regolith at approximately 2110 kg/m3 [3].
The assumptions made above about the type of regolith material found on the surface of Ceres
drive key aspects for the ISPP facility. By determining the specific heat of the regolith, we can
estimate the power required to heat the rocks to extract the water trapped inside. The specific
heat of the regolith is estimated by using the following equation [4]:
(4.1.2-1)
We assume that the density of the regolith is similar to the density of the moon rocks found in
the original Apollo missions [3]. We also assume the porosity of the material, , to be large
enough to hold the 3% of water the regolith contains (as given in our Mission Design
Requirements). These assumptions lead to the specific heat of the regolith being 0.84 J/gK. To
Project Conclusions General Project Concerns Page 73
Author: David Wyant
Co-Author: Zachary Richardson
assure that the water is fully extracted from the regolith, another mission parameter requires
heating the regolith to 200oC above zero.
Project Conclusions General Project Concerns Page 74
Author: David Wyant
Co-Author: Zachary Richardson
References
[1]. Howard, K.A., "Lunar Avalanches," Abstracts of the Lunar and Planetary Science
Conference, Vol. 4, 1973, pp. 386.
[2]. Moore, H., Pike, R., and Ulrich, G., "Lunar Terrain and Traverse Data for Lunar Roving
Vehicle Design Study," Lunar and Planetary Institute, Houston, TX, March 1969.
[3]. Korotev, Randy L., (2004). "Density and Specific Gravity", Meteorites and Meteorwrongs,
Department of Earth and Planetary Sciences., Washington University of St. Louis. Date
Accessed: Feb. 15, 2011. URL: http://meteorites.wustl.edu/id/density.htm.
[4]. Waples, Douglas W., Waples, Jacob S., "A Review and Evaluation of Specific Heat
Capacities of Rocks, Minerals, and Subsurface Fluids. Part 2: Fluids and Porous Rocks".
Natural Resources Research, Vol. 13, No. 2, June 2004.
[5] Carrier, III, W. David. "Geotechnical Properties of Lunar Soil." Lunar Geotechnical Institute
(2005): 1-24. Web. March 6, 2011.
Project Conclusions General Project Concerns Page 75
Author: Jillian Roberts
4.1.3. Consideration of Space Suits
Our mission is designed to have a “shirt-sleeve” working environment, meaning that, in
nominal operation, astronauts should never have to don a spacesuit to complete the mission.
Only during launch, landings, and in emergency situations will an astronaut wear a spacesuit.
Should such a contingency arise, the spacesuit should be lightweight and flexible to reduce
fatigue during an extravehicular activity (EVA). The gloves must preserve as much dexterity as
possible for working with tools. The suit must protect the astronaut from the harsh environments
encountered on Ceres or in space.
We protect the astronauts in emergency situations using the Bio-Suit, developed at
Massachusetts Institute of Technology (MIT) and currently at TRL 4. The Bio-Suit is a modular
design based on mechanical counter pressure, using elastic tension rather than gas pressurization.
It decreases the risk for depressurization, and allows greater freedom of movement [1].
When a human moves, the skin stretches and compresses with the motion. However, there
are certain lines on the skin that do not deform, called “lines of non-elongation”. The suit takes
advantage of this skin property by orienting elastic mesh fibers along lines of non-elongation and
maximizing mobility. In essence, the suit is truly a second skin [2]. Figure 4.1.3-1 shows elastic
cords along the lines of non-extension [3] and the prototype Bio-Suit elastic skin, worn by Prof.
Dava J. Newman (Photo copyright of Volker Steger/Science Photo Library) [4].
Project Conclusions General Project Concerns Page 76
Author: Jillian Roberts
Figure 4.1.3 Elastic cords follow the lines of non-elongation [3], and Prof. Deva J. Newman
demonstrates the suit’s flexibility by jumping [4].
We gain another factor of reliability through the Bio-Suit layers, which can locally self-repair
and preserve integrity of the suit. If a small hole does appear, the suit does not lose breathable
oxygen and the astronaut‟s skin remains unharmed. If a hole larger than 1 mm2 appears, the
astronaut would have time to return to a safe environment before the reduced pressure causes
significant damage to his skin [1].
By: Volker Steger By: Dava J. Newman
Project Conclusions General Project Concerns Page 77
Author: Jillian Roberts
Figure 4.1.3-2 The modular Bio-Suit design is easy to don [3].
Each suit is custom fitted with laser scanning and an electrospinlacing process, ensuring
proper tension in the suit skin. Remaining components on the suit are simple, interchangeable,
and easily maintained. Figure 4.1.3-2 above shows the simplicity of donning the modular
spacesuit [3]. Astronauts can tweak the elastic suit size real-time, accommodating changes such
as muscle atrophy, weight gain or loss, and spinal elongation. The full suit consists of the
elastic Bio-Suit layer and a hard torso shell with portable life support, which provides gas
counter pressure [1].
Table 4.1.3-1 The Bio-Suits for the entire crew have specifications as shown below.
Mass, kg Power, kW Volume, m3
Bio-Suits 216 0.0176 0.0432
Because the suit is at a Technology Readiness Level of only about 4, specifications like mass,
power, and volume are not well publicized. We use several approximations to calculate these
By: Dava J. Newman
Project Conclusions General Project Concerns Page 78
Author: Jillian Roberts
values, the details of which are in the Appendix. The table above shows the total mass, power,
and volume for Bio-Suits for a 6-member crew.
To mitigate risks, the astronauts will wear Bio-Suits during the mission phases with the
highest likelihood of failure occurrence - during launch, landing on Ceres, moving between
stations, takeoff from Ceres, and Earth re-entry. The astronauts will be able to survive an abort
during launch or sudden cabin depressurization during any of these phases.
Project Conclusions General Project Concerns Page 79
Author: Jillian Roberts
References
[1] Pitts, Bradley, et al. Astronaut Bio-Suit for Exploration Class Missions: NIAC Phase I
Report, 2001.Massachusetts Institute of Technology, 2001.
[2] Newman, Dava J. “Bio-Suit Patterning: Testing the Line of Non-Extension”. Astronaut
Bio-Suit System for Exploration Class Missions. Bimonthly Report. Massachusetts
Institute of Technology, March 2005.
[3] Newman, Dava J. (2004, April). An Astronaut „Bio-Suit‟ System for Exploration
Missions. Presented at workshop in Massachusetts Institute of Technology in Cambridge,
Massachusetts.
[4] Extra-Vehicular Activity (EVA) Research @ MVL, BioSuit –Overview. [Retrieved] 24
Feb 2011. [from] http://mvl.mit.edu/EVA/biosuit/index.html
Project Conclusions Detailed Mission Timeline Page 80
Author: Courtney McManus
Co-Author: Graham Johnson
4.2. Detailed Mission Timeline
Our mission is designed to span just under eleven years from the first launch to crew
splashdown, with a total crew mission time of 3.69 years. We identify a total of thirteen mission
phases which are shown in the Fig. 4.2-1 below. Each of the phases of the mission is described in
the following paragraphs.
Figure 4.2-1 Graphic depiction of the mission chronology from first launch to splashdown
The timeline given in the following section is just one possible timeline our mission could
follow. We chose to present this timeline because it illustrates the earliest-possible time frame
for the mission.
Phase 1 - STV Launches and Construction
The first launch of Project Vision takes place on January 1st, 2020. This launch carries
elements of the Supply Transfer Vehicles to low Earth orbit to begin construction. We follow
this launch with approximately one launch every two months for a total of 19 launches. The total
construction time for the two STVs is three years.
Project Conclusions Detailed Mission Timeline Page 81
Author: Courtney McManus
Co-Author: Graham Johnson
Phase 2 – STV Transfer to Ceres
On October 7th
, 2024, both Supply Transfer Vehicles begin the interplanetary transfer to
Ceres. Both vehicles begin transfer on the same day, and these transfers last approximately 1.4
years.
Phase 3- Cargo Deployment and ISPP Production
STV 1 reaches orbit around Ceres on March 6th
, 2025; STV 2 arrives on March 27th
, 2025.
From here, we land the ISPP facilities and rovers at their appropriate places on the surface. At
this time, we set up and calibrate the facilities to begin propellant and consumables production.
The ISPP facilities take 2.25 years to produce the requisite propellant, with ISPP Station
1finishing the process on June 10th
, 2027, and ISPP Station 2 finishing on June 30th
2027.
Phase 4 – Halo Satellite Transfer to Orbits
After the STVs reach orbit around Ceres, we release the halo satellites to being their journey
to their respective Lagrange points on March 27th
, 2025. This transfer takes the satellites 1.8
years to accomplish, arriving in their proper orientation on February 8th
, 2027. These transfers
happen in concurrence with the ISPP production, and we note that, during their transfers, the
halo satellites are in a useable configuration.
Phase 5 – CTV Launches and Construction
We launch the first components of the CTV on February 11, 2027. Again, we assume that we
are able to launch one Ares V every two months, which gives a construction time of about 1.5
years.
Phase 6 – Crew Launch
We launch the crew in the Crew Capsule atop an Ares I rocket on August 18th
, 2028. The
Capsule then rendezvous with the CTV and the crew ingresses to their new home in space.
Project Conclusions Detailed Mission Timeline Page 82
Author: Courtney McManus
Co-Author: Graham Johnson
Phase 7 – CTV Transfer to Ceres
We begin the interplanetary transfer of the CTV on August 19th
, 2028. This transfer lasts 1.38
years, with the CTV arriving in a low Ceres orbit on January 17th
, 2030.
Phases 8 & 9 – Exploration at ISPP Stations
The CTV lands at ISPP Station 1 and the crew begins exploration of the planet in the rovers.
For 196 days, the crew stays at the first station, after which they transfer to ISPP Station 2. The
crew remains here for another 196 days before launching to a low Ceres orbit. The crew is on the
surface of Ceres for a total of 392 days.
Phase 10 – CTV Transfer to Earth
On February 13th
, 2031, the CTV leaves low Ceres orbit and begins the interplanetary transfer
back to Earth. This transfer takes 1.25 years.
Phase 11 – CTV and Capsule Aerocapture
The CTV enters the first boundaries of the Earth‟s atmosphere on May 12th
, 2031. At this
time, the crew ingresses into the Crew Capsule which is then separated from the CTV. Each
vehicle then opens a ballute to aerocapture with the Earth‟s atmosphere. This initial aerocapture
can last from less than 1 day to about a week, depending on the density of the atmosphere at the
time and position of capture.
Phase 12 – Capsule Re-entry and Splashdown
After the aerobrake maneuver is complete, the Crew Capsule jettisons the ballute and begins a
controlled descent through the Earth‟s atmosphere. The Capsule uses three parachutes to slow
the descent to an eventual splashdown in the ocean.
Project Conclusions Detailed Mission Timeline Page 83
Author: Courtney McManus
Co-Author: Graham Johnson
Phase 13 – CTV Aerobraking to Stable Low Earth Orbit
After the Capsule is jettisoned, the CTV continues to use the ballute to slow itself down to a
stable low Earth orbit. The CTV remains here in a reusable configuration for future mission use.
Project Conclusions Estimated Mission Cost Page 84
Author: Courtney McManus
4.3. Estimated Mission Cost
Space travel is expensive. A typical cost for today‟s interplanetary missions can range from
$1,500 million to $3,000 million – and such a price is for a single satellite like Cassini or
Galileo, not something as complex and massive as our Crew Transfer Vehicle [1]. So, it stands to
reason that our mission will be very expensive. Because of the high cost of our project, dollar
values mentioned here will be in billions of US dollars circa the year 2011(B USD „11), unless
otherwise noted.
To estimate the total cost of our mission, we employ the Advanced Missions Cost Model
(AMCM) developed by NASA Johnson Space Center [2]. This model takes information from a
database of over 260 past space missions to provide a cost estimate based on the following
criteria:
1. Quantity, Q: The number of units produced, including production units, spares, and
redundancies
2. Mass, M: The dry mass of the system, in pounds
3. Specification, S: A designator to the type of mission flow (for example,
communications, planetary lander, launch vehicle upper stage, etc.)
4. Initial Operational Capacity, IOC: This is the first year which the system will be used
in operation
5. Block, B: A designator which specifies the level of design inheritance
6. Difficulty, D: Assess the difficulty of developing and producing the element
For more information on ACMC, please see section A4.3 of the accompanying appendix and the
resources listed in the reference section. In conjunction with the AMCM, we estimate the cost
of launching the components to low Earth orbit on Ares V and Ares I vehicles.
Project Conclusions Estimated Mission Cost Page 85
Author: Courtney McManus
4.3.1 Vehicle Costs
Crew Transfer Vehicle
We create a single Crew Transfer Vehicle for this mission, which has a total mass of
approximately 170,000 kg. (We note that the AMCM was computed using pounds for the
masses, however the information here will be presented in kilograms.) The CTV is a new design,
so we set the Block number to be one and the Difficulty to be high. We begin operational use of
this Manned Habitat in the year 2027. The following table highlights these values and shows the
total cost of the CTV.
Table 4.3.1-1: CTV Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 1
Dry Mass kg 170,000
Specification - - Manned Habitat
Initial Operating Capability - - 2027
Block - - 1
Difficulty - - High
Cost
Billion USD ’11 $23.39
Supply Transfer Vehicles
We have two Supply Transfer Vehicles for this mission, the bigger with a mass of 62,000 kg.
We use this bigger mass for a more conservative estimate of the cost. The STV design is a new
one, so we set the Block number to one and the Difficulty to high. These Planetary spacecrafts
will begin operation in 2020. The following table highlights these values and shows the total cost
of both STVs.
Project Conclusions Estimated Mission Cost Page 86
Author: Courtney McManus
Table 4.3.1-2: STV Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 1
Dry Mass kg 62,000
Specification - - Planetary
Initial Operating Capability - - 2020
Block - - 1
Difficulty - - High
Cost
Billion USD ’11 $53.47
In-Situ Propellant Production Stations
We have two ISPP Stations for this mission, each with a mass of around 28,000 kg. These
Stations are a new design, so we set the Block number to one. Since the system is quite complex,
we set the Difficulty to very high. These Planetary Landers begin operation in 2020. The
following table highlights these values and shows the total cost of both ISPP Stations.
Table 4.3.1-3: ISPP Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 2
Dry Mass kg 28,000
Specification - - Planetary Lander
Initial Operating Capability - - 2020
Block - - 1
Difficulty - - Very High
Cost
Billion USD ’11 $67.01
Exploration Rovers
For this mission, we use to Exploration Rovers to explore the surface of Ceres. Each rover has
a mass of approximately 11,510 kg. The Rovers have some heritage technology and design, so
we set the Block number to two. The Rovers leave Earth in 2020 and have an average Difficulty
Project Conclusions Estimated Mission Cost Page 87
Author: Courtney McManus
level. The following table highlights these values and shows the total cost of both Exploration
Rovers.
Table 4.3.1-4: Exploration Rovers Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 2
Dry Mass kg 11,510
Specification - - Rover
Initial Operating Capability - - 2020
Block - - 2
Difficulty - - Average
Cost
Billion USD ’11 $3.02
Rescue Rover
Our single Rescue Rover has a mass of approximately 6,500 kg and is the first of its kind, so
we set the Block number to one. The Rover begins operation in 2020 and has a high Difficulty.
The following table highlights these values and shows the total cost of the Rescue Rover.
Table 4.3.1-5: Rescue Rovers Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 1
Dry Mass kg 6,200
Specification - - Rover
Initial Operating Capability - - 2020
Block - - 1
Difficulty - - High
Cost
Billion USD ’11 $2.74
Ceres Orbiting Communication Satellites
We place two communications satellites in halo orbits around Ceres, each with a mass of
approximately 13,000 kg. These Communication satellites have some heritage technology, so we
Project Conclusions Estimated Mission Cost Page 88
Author: Courtney McManus
assign them a Block number of two and a Difficulty of average. They leave the Earth in the year
2020. The following table highlights these values and shows the total cost of the two Halo
Satellites.
Table 4.3.1-6 Halo Orbiting Communication Satellites
Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 2
Dry Mass kg 13,000
Specification - - Communication
Initial Operating Capability - - 2020
Block - - 2
Difficulty - - Average
Cost
Billion USD ’11 $4.62
Earth Trailing Relay Satellite
We place one Communications satellite in an earth-trailing heliocentric orbit to complete our
communications network. This satellite has a mass of roughly 8,800 kg and involves heritage
technology, so we assign it a Block number of two. The satellite will begin operation around
2020 and has a Difficulty level of low. The following table highlights these values and shows
the total cost of the Relay Satellite.
Table 4.3.1-7 Relay Communication Satellite
Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 1
Dry Mass kg 8,800
Specification - - Communication
Initial Operating Capability - - 2020
Block - - 2
Difficulty - - Low
Cost
Billion USD ’11 $0.91
Project Conclusions Estimated Mission Cost Page 89
Author: Courtney McManus
Crew Capsule
The Crew Capsule takes the crew to and from low Earth orbit and the Earth‟s surface, and has
a mass of approximately 10,000 kg. We classify this manned re-entry vehicle as having a Block
number of two, with a Difficulty of average. The Capsule will begin operation in the year 2027.
The following table highlights these values and shows the total cost of the Crew Capsule for our
mission.
Table 4.3.1-7 Crew Capsule Advanced Mission Cost Model breakdown
Input Parameter Units Value
Quantity units made 1
Dry Mass kg 10,000
Specification - - Manned Re-entry
Initial Operating Capability - - 2027
Block - - 2
Difficulty - - Average
Cost
Billion USD ’11 $4.10
Project Conclusions Estimated Mission Cost Page 90
Author: Courtney McManus
4.3.2 Launch Costs
In order to successfully model a total cost estimate of our mission, we need to also account for
the cost of the launch vehicle operations. We estimate the approximate cost per launch of the
Ares V vehicle to be $500 million USD [3]. Similarly, we estimate the cost of an Ares I launch
to be approximately $150 million USD [4]. The following table shows the number of launches
we require, as well as the total launch cost for our mission.
Table 4.3.2-1 Launch costs for Project Vision
Launch Vehicle Number of Launches
Needed
Cost, Billion USD ‟11
Ares V 26 $13
Ares I 1 $0.15
Launch Cost
Billion USD ’11 $13.15
Project Conclusions Estimated Mission Cost Page 91
Author: Courtney McManus
4.3.3 Overall Mission Cost Estimate
We total the total mission cost estimate by combining the estimated cost of each vehicle with
the estimated costs of the Ares V and Ares I launches. We total these costs in the following table.
Table 4.3.3-1 Total estimated mission cost
Mission Element Element Cost,
Billion USD ‟11
Crew Transfer Vehicle $23.3
Supply Transfer Vehicles $53.5
ISPP Stations $67.0
Exploration Rovers $3.0
Rescue Rover $2.7
Ceres Orbiting Comm. Sat. $4.6
Relay Comm. Sat. $0.9
Crew Capsule $4.1
Ares V Launches $13.0
Ares I Launches $0.15
Total Mission Cost Estimate
(Billion USD ’11) $172.70
With this total mission cost estimate, we find some interesting breakdowns of the cost. By
dividing the total cost of the span of the mission, we find the total cost per year of our mission to
be $15.66 billion (USD ‟11) each year. We find that the overall cost per kilogram of dry mass of
our mission is $656 million (USD ‟11) per kilogram. These values are shown in the following
table:
Table 4.3.3-2 Cost parameters of Project Vision
Cost Parameter Cost,
Billion USD ‟11
Total Mission Cost $172.70
Cost per year $15.66 per year
Cost per kilogram dry mass $0.66 per kg
Project Conclusions Estimated Mission Cost Page 92
Author: Courtney McManus
4.3.4 Cost Comparisons
The closest program with which we can compare our total estimated mission cost is the
Apollo Program in the 1960s and 1970s. The total cost of the Apollo Program is estimated to be
around $20 billion (USD ‟65) [5]. We compensate this amount for inflation and find that this
equates to roughly $140 billion (USD ‟11) in today‟s economy [6]. As stated before, our total
estimated mission cost is roughly $173 billion (USD ‟11). From this estimate, we find that our
mission has an estimated cost approximately 20% higher than that of the Apollo Program. This
stands to reason when we consider the complexity of the mission, the distance of Ceres, the
technological advances that will need to be made, the duration of the mission, and other
parameters.
Project Conclusions Estimated Mission Cost Page 93
Author: Courtney McManus
4.3.5 Possible Means of Cost Reduction
In retrospection, we identify a few changes that could be made to the mission architecture
which might reduce the cost of the overall mission.
The first change that we could make is to use an engine other than chemical engines for the
kick motors on the STV and CTV. These engines require massive amounts of propellant to be
sent to LEO on the Ares V. Reducing the number of launches of Ares V needed to fuel the
engines would reduce the overall cost of the mission.
We find that most methods for cost reduction have an impact in the risk of a loss of crew
catastrophe. We can make these tradeoffs on a case-by-case basis in the future to reduce costs.
Such options include having only one ISPP station on Ceres, not toting the Crew Capsule all the
way to Ceres on the CTV, and increasing the time of interplanetary transfer. For more
information on the Risk Analysis completed for this mission, please see section 4.4 of this report.
Project Conclusions Estimated Mission Cost Page 94
Author: Courtney McManus
References
[1] Wertz, James R. (ed) and Larson, Wiley J. (ed), Reducing Space Mission Cost, Microcosm
Press, El Segundo, CA, 1996.
[2] Larson, Wiley J. (ed), and Pranke, Linda K. (ed), Human Spaceflight Mission Analysis and
Design, The McGraw-Hill Companies, New York, 1999.
[3] Stahl, Philip H., “Ares V Launch Capability Enables Future Space Telescopes,”
International Society for Optics and Photonics Conference, 6687-16, 2007.
[4] Smith, Marcia, “How Much Would Ares I Cost?,” Space Policy Online, March 2010.
[http://spacepolicyonline.com/pages/index.php?option=com_content&view=article&id=81
7:how-much-would-ares-i-cost&catid=67:news&Itemid=27]
[5] Launius, Roger D., “Project Apollo: A Retrospective Analysis,” NASA Office of Policy
and Plans, History Office, 2004.
Resources
- US Inflation Calculator, CoinNews website family, accessed March 2011.
[http://www.usinflationcalculator.com/]
- Advanced Mission Cost Model Online Tool, NASA Cost Estimating Website.
Accessed March 2011. [http://cost.jsc.nasa.gov/AMCM.html]
Project Conclusions Risk Assessment Page 95
Author: Courtney McManus
4.4. Risk Assessment
Spaceflight is a risky business. Although it is difficult to say with exact precision the
probability that a mission will succeed or fail, we follow the NASA Exploration Systems
Architecture study [1] to get a good idea of the probability of our mission succeeding. It is
important to note that this risk assessment is just a preliminary study and obtaining more
definitive conclusions will require a more in-depth study. For this project, we concern ourselves
with only the Loss of Crew risk probabilities, as this is the only parameter specified in the
Mission Design Requirements. As such, any Loss of Mission event occurring before the crew is
launched from Earth is not taken into account in this risk analysis. We show the risk probabilities
of Project Vision in the following table.
Table 4.4-1 Project Vision risk assessment probabilities
Event Failure % Success %
Launch to LEO 0.1 99.9
Capsule repositioning 1.1 98.9
Transfer to Ceres 0.3 99.7
Landing on Ceres 0.1 99.9
Quiescent Ops on Ceres 0.7 99.3
ISPP Failure 0.3 99.7
CTV transfer to Station 2 0.2 99.8
Launch from Ceres 0.2 99.8
Transfer to Earth 0.3 99.7
Aerocapture 1.2 98.8
Re-entry 0.1 99.9
Solar Particle Event 3.75 96.25
Total Probability of Safe Crew Return 91.6%
A more detailed explanation of each of the Events shown in Table 4.4-1 is found in section
A4.4 of Appendix A.
Project Conclusions Risk Assessment Page 96
Author: Courtney McManus
4.4.1 Possible Ways to Improve Risk Probabilities
We note that our probability of success is about 91.6%, which is below the Mission Design
Requirement of 95%. We identify a few possible ways in which we could reduce the overall risk
of our mission, at a tradeoff of cost.
Looking at Tale 4.4-1, we see that the biggest single source of risk for Project Vision is the
possibility of a Solar Particle Event (SPE). For every year the crew spends in space we subtract
one percent from our success percentage to account to the possibility of a 100-year, catastrophic
release of energy from our sun. We could reduce the amount of time the crew spends in space by
decreasing the duration of the transfers to and from Ceres. To accomplish this, we would have a
bigger change of velocity kick (ΔV) on the Crew Transfer Vehicles. Although this change would
reduce the risk, the overall cost of the mission would increase due to the increase in the mass of
propellant needed for the kicks.
Another significant source of risk for our mission is the repositioning of the Crew Capsule to
different locations throughout the mission. The decision to take the Crew Capsule with us to
Ceres was made based on initial analysis showing that such a configuration would return the
crew back to Earth faster once aerobraking had begun. If the Crew Capsule was not taken to
Ceres, but instead rendezvoused with the CTV upon return to the Earth, the risk of the multiple
repositionings of the Capsule would be eliminated.
Another way we identify to reduce the risk of the mission is to place the ISPP Stations next to
one another, as opposed to at antipodes of Ceres. This placement would eliminate the need to
transfer the CTV from the first Station to the second, while still maintaining the necessary
redundancy of having two stations.
Project Conclusions Risk Assessment Page 97
Author: Courtney McManus
4.4.2 Contingencies and Redundancies
In order to help reduce the risk of our mission, we place certain redundancies and
contingencies within the designs of our vehicles. More information on these contingencies can be
found in the detailed vehicle descriptions of each vehicle (Section 5).
CTV Contingencies
We design the tethers with which we extend the Crew Transfer Vehicle to have a factor of
safety of 10. We ensure that each onboard system has sufficient back-up and redundant
equipment such as computers and controllers. Within the CTV habitat, we provide the crew with
ample radiation shielding, including a safe room which has more shielding that the rest of the
habitat in case of a predictable solar particle event.
Crew Launch Contingencies
We launch the crew to the CTV atop an Ares I rocket, with a Launch Abort System
incorporated into the Capsule interfaces. In the event of an emergency on the Launch or in the
first phases of the launch, the Launch Abort System fires, pulling the crew to safety from the
rocket stack.
We rendezvous the Crew Capsule with the CTV in low Earth orbit. If a failure were to occur
during the rendezvous and proximity operations, the crew is able to de-orbit and return safely to
Earth.
STV Contingencies
We back up the electromagnets holding the cargo components on the STV with mechanical
connections. Should the electromagnets fail for some reason these physical connections will
prevent the components from drifting apart during transfer.
Project Conclusions Risk Assessment Page 98
Author: Courtney McManus
Exploration Contingencies
There are many contingencies worth mentioning for the exploration of Ceres. We design each
Rover to have two docking ports which can connect to either of the two docking ports on the
CTV. In the event that one port should fail for any reason, the others can be used. We include
spacesuits inside each of the Rovers for the crew to use should we lose atmospheric control. We
design a Rescue Rover to be used solely as an emergency response system for the two
Exploration Rovers. Each of the three Rovers is equipped with attitude and control correction
plans, should the Rovers flip over or be in an un-useable orientation.
ISPP Contingencies
We place two ISPP Station on the surface of Ceres to act as redundancies for each other. We
ensure that each Station is fully filled with enough propellant and consumables for the duration
of the surface stay, as well as for the return journey to Earth before the crew launches to the
CTV.
Crew Return Contingencies
We use three parachutes during the re-entry and splashdown phase of the Crew Capsule‟s life.
The system is designed to function properly on two parachutes. This means if one of the
parachutes should for some reason fail, the crew will experience no ill effects.
Project Conclusions Risk Assessment Page 99
Author: Courtney McManus
References
[1] Cirillo, W.M., Letchworth, J.F., Putney, B.F., Fragola, J.R., Lim, E.Y., Stromgren, C.,
“Risk-Based Evaluation of Exploration Architectures,” NASA Exploration Systems
Architecture Study Section 8, January 2005.
Project Conclusions Closing Comments Page 100
Authors: Courtney McManus, Evan Helmeid
4.5. Closing Comments
This semester-long feasibility study investigates the components necessary to send six people
to the dwarf planet Ceres and back. To accomplish this mission, we use twelve different vehicles
including Supply Transfer Vehicles (Cassiopeia and Cepheus), a Crew Transfer Vehicle
(Damocles), a Crew Capsule (ARC), two In-Situ Propellant Production facilities (APES 1&2),
two Exploration Rovers (Castor and Pollux), one Rescue Rover (SPRINT), and a
communications network consisting of three satellites (ECCO 1, 2, & Base).
We complete our mission with the goal of scientific exploration in mind. As such, we equip
our rovers and Crew Transfer Vehicle with many instruments of scientific study in the fields of
astronomy, geology, physics, chemistry, etc. Since ours will be a long-duration, crewed mission,
we will also be able to note the effects of prolonged spaceflight and limited-gravity environments
on human beings.
We conclude that our mission to Ceres is feasible, provided the assumed technological
advancements are made and that we have the necessary political and funding support. The
mission architecture we present follows the Mission Design Requirements laid out at the
beginning of the semester, has an overall cost of $173 Billion (USD ‟11), and has a probability
of success of 91.6%.
Detailed Vehicle Descriptions Page 101
5. Detailed Vehicle Descriptions
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 102
Author: Jared Dietrich
5.1. Supply Launch Vehicle
The Supply Launch Vehicles are a series of Ares V’s that we will use to move the Supply
Transfer Vehicle components into low Earth orbit. We select the Ares V due to its large volume
and mass capacity, and each payload has been designed to fit into the Ares V payload bay.
5.1.1 Launch Vehicle Selection
To determine Team Vision’s Supply Launch Vehicle (SLV), we employ a trade study of
current and proposed launch vehicles. The proposed system must meet a technology readiness
level (TRL) 3 for consideration. TRL 3 is defined as having “Analytical and experimental
critical function and/or characteristic proof of concept” [1]. We consider only relevant systems;
a vehicle with the capabilities of meeting our demanding payload mass and volume
requirements. Finally, we consider the cost of launching our massive payloads into Low Earth
Orbit (LEO).
The characteristics of several SLV options are found in table 1. Vehicles chosen are some of
the largest available and met our preliminary requirements. The trade study includes Ares I,
Ares V, Falcon 9, and Atlas V launch vehicles. From the information collected, we see great
benefits from the Ares V over all other options. The Ares V’s payload mass and vehicle
diameter are the largest of any available launch vehicle and meets the required TRL 3
requirement due to NASA’s extensive testing for the Constellation Program [2].
Table 5.1-1 SLV trade study
Launch Vehicle Height, m Diameter, m Mass, kg Thrust, kN Payload, kg
Ares I 94 5.5 927,142 17,180 25,400
Ares V 116 10 3,704,534 32,629 188,000
Falcon 9 54.3 3.66 885,000 4,940 32,000
Atlas V 59.7 3.81 565,768 8,590 29,420
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 103
Author: Jared Dietrich
Although the Ares V is vastly superior in performance, we must also consider a cost
comparison between it and the competition. Figure 5.1-1 below shows the cost per kilogram for
all four launch vehicles in our study, where we normalize launch cost by each vehicle’s payload
mass delivered into LEO.
Fig. 5.1-1 Launch cost per payload mass delivered into LEO
Ares V offers the least cost to deliver payload into LEO at $1,826/kg. Therefore, Ares V is
our choice for Project Vision’s SLV. The Ares V delivers up to 188 T to LEO and has a usable
payload volume of 1,410 m3 [3]. Although the Ares V has yet to be launched, it employs reliable
technology and systems. The Ares V’s six RS-68B main engines are derived from the Space
Shuttle Main Engine (SSME) and Delta IV family of launch vehicles. In addition, two five and a
half segment Solid Rocket Boosters (SRBs) flank the Ares V core stage. These boosters are also
derived from the Space Shuttle’s SRBs. With tried and tested technology, the Ares V offers high
performance and high reliability.
Ares 1 Falcon 9
Atlas V
Ares VFalcon 9 (H)
Atlas V (H)
$0
$2,000
$4,000
$6,000
$8,000
$10,000
$12,000
$14,000
$16,000
1 2 3 4
Co
st/k
g
Launch Vehicle options
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 104
Author: Andrew Curtiss
Co-author: Megan Sanders
5.1.2. Launch Manifest and Timeline
We organize the launches so that we are able to provide power to the magnetic connections as
well as keep the Supply Transfer Vehicle mass as balanced as possible during construction.
Table 5.1.1-1 and Table 5.1.1-2 detail the launch order for the components of both Cassiopeia
(STV 1) and Cepheus (STV 2).
Table 5.1.2-1 Launch Manifest for STV 1
Launch Number Cargo Total Mass, kg Total Volume, m3
1 Center Module 124446.96 585.73
2 LH2 Tank 1 84648.00 1164.70
3 LH2 Tank 2 84648.00 1164.70
4 LOX Tank 1 146920.00 181.29
5 LOX Tank 2 146920.00 181.29
Low Thrust Engine 664 0.005
6 LOX Tank 3 146920.00 181.29
7 LOX Tank 4 146920.00 181.29
Low Thrust Engine 664 0.005
8 LOX Tank 5 146920.00 181.29
Low Thrust Engine 664 0.005
9 Reactor 22926.27 296.00
Table 5.1.2-2 Launch Manifest for STV 2
Launch Number Cargo Total Mass, kg Total Volume, m3
10 Center Module 117799.27 412.05
11 LH2 Tank 1 84567.00 1164.70
12 LH2 Tank 2 84567.00 1164.70
13 LOX Tank 1 145850.00 181.29
14 LOX Tank 2 145850.00 181.29
Low Thrust Engine 664 0.005
15 LOX Tank 3 145850.00 181.29
16 LOX Tank 4 145850.00 181.29
Low Thrust Engine 664 0.005
17 LOX Tank 5 145850.00 181.29
Low Thrust Engine 664 0.005
18 Reactor 22926.27 296.00
19 Comm. Satellites 28347.81 463.53
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 105
Author: Andrew Curtiss
Co-author: Megan Sanders
The center modules contain all of the payload components for each of the vehicles, attitude
control systems, telemetry reporting communication systems, rocket engines, module connectors,
and the landing gear. Table 5.1.1-3 and Table 5.1.1-4 contain manifests of the payload and
structural components in each of the center modules. The structural components are outlined in
detail in the Supply Transfer Vehicle sections.
Table 5.1.2-3 STV 1 Center Module Payload Contents
Cargo Mass, kg Volume, m3 Quantity Notes
Exploration Rovers 23005.11 131.484 2
Rescue Rover 6,413.32 42.13 1
Food 8164.5 53.641 3.5 Years
Radiation Shielding 23912 5.8324 1 box
Module Connectors 733.26 1.26 24
Landing Legs 275.8344 0.1388 4
Thermal Control System 1,060.58 1.864 --
Telemetry Dish 1.7 0.005 --
Computers 29.03 0.025 --
Kick Engines 7967.64 73.2 3
Payload Storage Container 5539.1 3.1118 1
Attitude Control Propellant 1279.76 1.064 --
Attitude Determination Hardware 10 0.005 --
Attitude Control Hardware 150 0.218 6
Landing Control Propellant 50 0.042 --
Antenna Pointing Controller 5 0.125 Canfield Joint
Kick Control Propellant 850 0.84 --
Ceres Regime Engines 181.17 0.07 1
Interstage 3023.3 61.56 --
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Co-author: Megan Sanders
Table 5.1.2-4 STV 2 Center Module Payload Contents
Cargo Mass, kg Volume, m3 Quantity Notes
Food 8164.5 53.641 3.5 Years
Radiation Shielding 23912 5.8324 1 box
Module Connectors 733.26 1.26 24
Landing Legs 275.8344 0.1388 4
Thermal Control System 1,060.58 1.864 --
Telemetry Dish 1.7 0.005 --
Computers 29.03 0.025 --
Kick Engines 7967.64 73.2 3
Payload Storage Container 5539.1 3.1118 1
Attitude Control Propellant 1532.98 1.277 --
Attitude Determination Hardware 150 0.218 6
Landing Control Propellant 50 0.042 --
Antenna Pointing Controller 5 0.125 --
Kick Control Propellant 825 0.840 --
Ceres Regime Engines 181.17 0.07 1
Interstage 3023.3 61.56 --
Upon arrival in low Earth orbit, the payload shrouds are jettisoned. The first launch contains
the center module. The center module features a payload storage container, which acts as a
structural shell to contain the payload. The following launches jettison the payload shrouds to
reveal propellant tanks, reactors, and the communication satellites. For the launches that contain
the engines, the payload shroud will allow for the engine to be attached before launch, reducing
the amount of construction needed to be performed in low Earth orbit.
We will launch the first vehicle on January 1, 2020 and proceed with a launch every 2
months. The launches and construction together will take a approximately three years and will
be completed on January 1, 2023. Due to the extended time needed for construction the launch
vehicle will also bring extra attitude control propellant that can be used to keep the transfer
vehicle stable. This extra propellant will be carried in small tanks attached to the outside of
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 107
Author: Andrew Curtiss
Co-author: Megan Sanders
launches 1 and 10. These tanks will be attached so that they can be separated before the supply
transfer vehicle departs.
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5.2 Supply Transfer Vehicle
5.2.1. Construction in LEO
The Supply Transfer Vehicle (STV) is made up of a center module with propellant tanks and
a power generating reactor. We designed the components in a flower pedal type arrangement
with the center module in the middle and propellant tanks and the reactor on the outside as
shown in Fig. 5.2.2-1. Each of the modules of the STV is unmanned and therefore requires
autonomous docking procedures in order to construct the both STVs in orbit.
We achieve rendezvous by docking two spacecraft (at a remote distance) together to become
one. In our case, the target spacecraft remains in a constant orbit while the other “chases” it
down until they rendezvous. More specifically, we refer to the assembled portion of the STV
(center module with or without some outer modules) as the chaser spacecraft and the newly
launched piece of the STV as the target. Only the chaser spacecraft performs orbital maneuvers
because only the center module features attitude hardware and propellant. When the target
spacecraft reaches LEO, the chaser spacecraft performs maneuvers to change its orbit and
orientation to rendezvous with the target spacecraft.
The spacecraft rendezvous with the help of Autonomous Orbit Control [1]. The Autonomous
Orbit Control system features software that tracks the location and orbital characteristics of the
target spacecraft at all times. The two spacecraft do not need to be communicating with each
other in order to perform the maneuvers. This simplifies the process significantly because we do
not require long-range communication between the spacecraft. The Autonomous Orbit Control
system adds no mass to the STV. The system uses existing hardware such as the computers
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onboard the STV to calculate trajectory operations that allow the chaser to reach the target
spacecraft’s orbit.
We launch each of the modules of the STV separately due to mass and volume constraints.
Consequently, we require each module to perform its own autonomous rendezvous procedure.
We launch the center module first and it becomes the first chaser spacecraft. We launch the
reactor second because its power is required to perform the rendezvous and docking procedures.
When the launch vehicle deposits the reactor payload into LEO, it becomes the first target
spacecraft. After the reactor reaches its constant orbit, the center module chaser calculates
trajectory and attitude operations and chases the target spacecraft. We repeat this process of
chasing down target spacecraft until the entire STV is assembled.
After the rendezvous maneuvers have placed the spacecraft next to each other, we must dock
them together. We perform the docking maneuver using the module connectors that physically
connect the spacecraft modules together. We power the electromagnets in the module connectors
on and off to capture and release modules. After the spacecraft achieves rendezvous, we turn on
the power to the magnets. Supplying power to the electromagnets not only creates a force that
binds the modules together, but also heats the connectors which causes the materials to expand.
This expansion tightens the bond between modules and provides a redundant docking
mechanism between modules in the event of a power failure.
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References
[1] Wertz, James R. and Bell, Robert, “Autonomous Rendezvous and Docking Technologies –
Status and Prospects” SPIE AeroSense Symposium Paper No. 5088-3, April 2003.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 112
Author: Andrew Curtiss
5.2.2. Configuration Overview
The Supply Transfer Vehicles consists of a center module, a reactor, and propellant tanks. The
center module features the payload storage container, kick and Ceres regime motors, a set of
landing legs, and the module connectors. Propellant tanks and the reactor surround the center
module in the flower pedal shape shown in Fig. 5.2.2-1.
Figure 5.2.2-1 Top view of the STV. Notice that the arrangement of the modules around the
center module is in a flower pedal formation.
LOX
Tanks
Reactor
LH2
Tank
LOX
Tanks
Center
Module
LOX
Tank
By: Andrew Curtiss
LH2
Tank
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Author: Sonia Teran
5.2.3 Trajectory
The goals that drove to our selection of the trajectories are:
1) All supplies must arrive to Ceres within 5 years of departing from Low Earth Orbit
(LEO) of 50 km
2) Provide the trajectory with the lowest propellant cost
In this feasibility study some key assumptions are used for designing these trajectories. The first
assumption is that the orbits are circular coplanar orbits. The second assumption is that the ΔV
maneuvers are considered to be impulsive maneuvers.
Our Supply Transfer Vehicles (STVs) begin their journey from LEO. In order to escape the
influence of Earth, they perform a ΔV maneuver. Keeping in mind our goal of low propellant
cost, we select the maneuver with the lowest ΔV required. The ΔV required for STV1 is 5 km/s
and for STV2 is 5.01 km/s. These trajectories by no means are optimal solutions; however, of all
the cases investigated these trajectories meet our goal. Determining the ΔV value can be found
in Section A.5.2.3 of the appendix.
We translate ΔV into mass of propellant (mpropellant) by rearranging the rocket equation as
follows,
(5.2.3-1)
where mpropellant is the mass of the propellant required, mwet is the mass of the vehicle with
propellant, ΔV is the change in velocity maneuver, Isp is the specific impulse of the engine, and
g0 is the reference gravity of Earth which is 9.80655 m/s2. The propellant required for this
maneuver is summarized in Tables 5.2.3-1 and 5.2.3-2. After this maneuver the STVs are now on
their heliocentric transfer to Ceres.
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We use a constant low thrust of 25 N throughout the entire heliocentric journey to Ceres for
both STVs. An algorithm with equations of motions to calculate the mpropellant requirement for the
low thrust trajectory calculates the position, velocity, and time of flight for the low thrust
trajectory. The state equations are as follows,
(5.2.3-2)
(5.2.3-3)
(5.2.3-4)
(5.2.3-5)
where r is the heliocentric position in km, θ is the angular position in radians, Vr is the radial
velocity in km/s, Vθ is the tangential velocity in radians/s, µ is the gravitational parameter of the
Sun in km3/s
2, T is the thrust, m is the mass of the vehicle determined by Eqn. 5.2.3-6, and α is
the steering law determined by Eqn. 5.2.3-7. The mass of the STVs are constantly decreasing
because we are always thrusting. The mass at any time can be calculated by
(5.2.3-6)
where m0 is the initial mass of the STV, is the mass flow rate, and dt is the time step. We
always thrust in the direction of our velocity and thus the steering law is,
(5.2.3-7)
Using these equations of motion and the steering law the heliocentric journey for STV1 takes
1.411 years and STV2 takes 1.466 years.
Once the STVs finish their heliocentric trajectory they are captured into a circular LCO of 50
km. They both arrive with excess velocity (V∞). STV1 arrives with a V∞ of 2.67 km/s excess
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 115
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velocity and STV2 arrives with 2.49 km/s. The ΔV required to capture each STV into a 50 km
LCO is 2.37 km/s and 2.19 km/s respectively. We calculate these values by the following
(5.2.3-8)
where μCeres is the gravitational parameter of Ceres and rLCO is the LCO altitude plus the radius of
Ceres. Using Eqn. 5.2.3-8 and Eqn. 5.2.3-1 we get the required propellant mass for the Ceres
capture maneuver. The trajectories of both vehicles are represented in Fig. 5.2.3-1. The red
portion of the figure represents the low thrust portion of the journey.
Figure 5.2.3-1 The general trajectory for both STV1 and STV2, where the red represents
the low thrust portion of the trajectory.
A breakdown of the propellant masses for each STV is summarized in Table 5.2.3-1 for STV1
and 5.2.3-2 for STV2.
By: Sonia Teran
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Author: Sonia Teran
Table 5.2.3-1 Propellant mass breakdown for STV1
Phase Propellant Mass, kg
at Earth 724,924
Low Thrust 22,695
at Ceres 134,773
Table 5.2.3-2 Propellant mass breakdown for STV2
Phase Propellant Mass, kg
at Earth 728,719
Low Thrust 23,575
at Ceres 127,240
In reality there is no system that can produce an impulsive maneuver, and a non-impulsive
maneuver requires more propellant. Therefore, a 15% propellant cost is added to all of the ΔV
maneuvers for burn arcs.
Criticism of Model
Using these trajectories with our assumptions we accomplish the goals for the STVs. Both
STVs arrive at Ceres under with more than enough time. As mentioned before we provide a non-
optimal trajectory solution for this problem. For a more accurate analysis, optimization
techniques should be used along with burn arcs and non-circular coplanar orbits. The assumption
for circular orbits is reasonable since the eccentricity of both bodies is nearly circular. The
Earth’s eccentricity is 0.0164 and Ceres’ is 0.079 [1]. However, the inclination of Ceres’ is
10.58º [1] from the ecliptic plane and will need to be taken into account for further analysis.
Also, to take advantage of the 5 year time of flight low thrust gravity assists could be used for the
STV trajectories.
The mass of propellant for the ΔV maneuvers can also be reduced if nuclear thermal engines
where used instead. However, for the purpose of this feasibility study and for comparison of
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known systems chemical rockets were chosen. A more detailed discussion can be found in
Appendix H.
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References
[1] “HORIZONS Web Interface,” Solar System Dynamics, URL:
http://ssd.jpl.nasa.gov/horizons.cgi#results [cited 26 March 2011].
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5.2.4. Power System
The design of the Supply Transfer Vehicle (STV) power system depends directly to the
proposed power budget of both STV 1 and STV 2. The biggest bulk of the power requirement is
determined by the power that is needed to operate the propulsion system for both STVs.
Table 5.2.4-1 STV propulsion power requirement
Vehicles Propulsion Power
Requirement, MW
STV 1 1.225
STV 2 1.225
We can see that the propulsion power requirements are the same for both STVs which led to
similar design for both of the power systems. The remainders of the power requirement are very
small compared to the power requirements of the propulsion system.
Table 5.2.4-2 STV auxiliary power requirements
Components Power Requirement , kW
Structures and Thermal 3.616
Communication 9.050
Attitude Control 0.080
TOTAL 12.746
With the proposed propulsion system, the power system consists mainly of an ultra-compact
high temperature molten sodium fast reactor combined with a thermo-photovoltaic (TPV) power
conversion system that powers both of the spacecraft with multiple megawatt capabilities. 2MW
of electrical power is a conservative desirable value for safety factors and assuming a decrease in
efficiency due to contamination and buildup of external particles and debris.
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A molten sodium fast reactor TPV power system components for each STV consists of a
reactor, radiation shielding, radiator, molten sodium, TPV conversion system, and armor.
Table 5.2.4-3 STV Power system components mass
Components Mass, kg
Reactor Core 235
Shielding 18008.7
Radiator 1393.33
Molten Sodium 875.72
TPV conversion system 100
Armor 1954.57
TOTAL 22288.5
Figure 5.2.4-1 General STV power system operating schematic
Molten Sodium
Armor
Radiant Heat
TPV Convertor
Reactor
Shielding
Radiator
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Figure 5.2.4-2 Model of the STV power system
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5.2.4.1 Reactor Design
The central and main component that provides thermal power that eventually converts into
electrical power is the reactor and the core. A safe and highly reactive core is desired since the
STV 1 and STV 2 need a minimum of 1.237 MW in full throttle for the transit the Ceres.
The main decision and limiting factors include:
- Small mass and volume
- Controllable
- Reliability and little maintenance required
In order to meet the requirements of small mass and volume, highly reactive fuels are
advantageous such as reactor grade plutonium carbide. Also, plutonium carbide has very good
thermal conductivity. For controllability, the reactor consists of Zr3Si2 rotatable reflectors [5] to
throttle how much heat can be transferred to the power conversion system and also for the uses
of shutting down the reactor during the end of life phase of the power system.
Liquid metal-cooled system pumps molten liquid through channels in the core to a power
conversion system to extract heat. This system is quite flexible and less massive than the other
reactor power system options available such as the gas cooled reactor and the heat pipe reactor.
Due to the high operating temperature of 1600K-1800K in a liquid metal-cooled system, the
radiator mass is considerably smaller than of direct-gas cooled system. Also due to its small
volume radiation shielding, the largest bulk of the power system mass, the total mass and volume
of the power system decreases considerable. Along with the liquid metal-cooled system, molten
salt is chosen as a heat transfer medium as the liquid due to its high heat transfer properties and
high boiling temperature. Due to liquid metal cooled system’s high operation temperature,
transfer medium (Molten Sodium) that has a high heat transfer properties and high boiling
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Author: Alex Park
temperature [13], and very compact overall mass and volume, we choose the molten sodium fast
reactor liquid metal cooled system to be the STV1 and STV2 reactor.
Table 5.4.2.1-1 Reactor specifications
Power provided 5MW thermal power
Reflector material Zr3Si2
Coolant / Heat transfer medium Molten Salt
Reactor Fuel Reactor Grade Pu (Plutonium) Carbide
Core dimensions Base Diameter=32 cm
Height=26 cm
Reactor/Fuel Mass 235 kg
Figure 5.4.2.1-1 Reactor depiction for the Supply Transfer Vehicle
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5.2.4.2 Thermo Photovoltaic (TPV) Cell Design
With limited or no human resources on the STVs, a power conversion system with little or no
maintenance required is very desirable. We choose the thermo photovoltaic (TPV) cells for the
power conversion system for the heat power provided by the reactor because of the
characteristics listed below.
1. High efficiency of 40%
2. No moving parts and lowers the maintenance required.
3. Simplistic and lightweight
4. Generates Direct Current (DC) electricity, which is mainly used by the electric propulsion
systems.
The operation of TPV cells is similar to that of solar cells. Instead of converting visible
light to electricity as in solar cells, TPV cells convert radiant heat energy generated by the
reactor. The biggest difference between the TPV cells and conventional photovoltaic cells is
that in TPV conversion system, the vehicle has control over the source instead of natural
sources such as the Sun.
The material that we choose for the TPV cells is Gallium Antimonide (GaSb) since they
the most logical choice for modern TPV generators for their efficiency and simplicity [11].
The GaSb TPV cell has an energy conversion efficiency of 44% [13]. We assume that the
conservative conversion efficiency is 40% due to safety factor and small malfunctions of the
TPV cells. The power generation density is 2 We/cm2 for the GbSb TPV cells [13].
The placement of the TPV cells is closely along the wall of the rocket armor. The TPV
cells start receiving thermal radiation from the internal radiators that transfers the hot molten
sodium in and out of the reactor. The idea of encasing the TPV cells in the rocket armor is to
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 125
Author: Alex Park
trap as much heat as possible to keep the operation temperature high and to reuse the
unconverted heat energy by trapping it inside the cylindrical case.
Table 5.2.4.2-1 TPV cell specifications
Material Gallium Antimonide (GbSb)
Power density 2 We/cm2
Conversion Efficiency 40%
Diameter 4.24 m
Height 15 m
Thickness 2 cm
Area 199.8 m2
Total mass 100 kg
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5.2.4.3 Radiator Design
In order for the TPV to convert thermal heat to electrical power, there has to be a good
source of radiant heat. The radiator is the source of the thermal heat energy that is transferred
from the hot molten sodium [10]. The radiator is a considerable part of the total mass of the
power system because it is dictated by the operating temperature.
(5.2.4.3- 1)
Ar = Total radiator area needed
= Stephan-Boltzmann Constant
= Emissivity
T = Operation temperature
Q = Heat power generated from the reactor
As we can see from the Equation 1 above, the mass of the radiator will decrease with the
fourth power of the operating temperature. The only way to keep the mass and volume of the
radiator small is to radiate the heat at a very high temperature.
In our case, the operating temperature of the molten sodium is 1600K and thermal energy
that needs to be radiated is 5MW. The conservative emissivity of the radiator is chosen as
0.85. The radiator area needed is 16.86 m2. The whole radiator system needs to be fit inside
the cylindrical power system with a 4.4 m diameter and a height of 15 m, the height of the
TPV conversion unit.
Shortest possible radiator piping is desirable since the operation temperature is very high
and the heat loss from the transit from the reactor to the radiator is very high. The best
geometry that would keep the radiator short and fits inside the cylindrical geometry of the
power system is to use a U shaped piping that would effectively circulate the molten sodium.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 127
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Figure 4.2.4.3-1The geometry of the piping dictates the mass of the molten sodium
coolant
kg
ddHm inout
MSMS 72.8754
22
(5.2.4.3- 2)
Where ρms is the Molten Sodium density, H is the total length, Dout is the outer diameter, and
Din is the inner diameter.
The calculation of the radiator mass results from the following equation.
kgHtddm Tiinoutradiator 33.1393 (5.2.4.3- 3)
Where Ti is the Titanium density, H is the total length, T is the thickness, Dout is the outer
diameter, and Din is the inner diameter.
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Table 5.2.4.3-1 Radiator Specification
Radiator Area 16.86 m2
Outer Diameter 0.224m
Inner Diameter 0.189m
Material Titanium
Radiator Mass 1393.33 kg
Molten Sodium Mass 875.72 kg
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5.2.4.4 Shielding Design
Two types of radiation, which are neutron and gamma radiation, emit from the reactor core
that contributes to the design factors and limitations of the radiation shielding of the STVs.
Neutron and gamma radiation from the core is damaging to the STV any biological materials
in the STV cargo and shielding is needed for the transit to Ceres and also when landing on
Ceres that when the crew arrives that the reactor radiation does not affect the human crew
members.
Certain material characteristics and properties are highly efficient as shielding materials.
High electron density per unit mass (Gamma radiation) and large neutron cross section per
unit mass (Neutron radiation) are two desirable attributes for a shielding material [7].
The shielding is coincidental to the reactor but below the reactor at the top of the STV
spacecraft, far away as possible from the cargo. We choose materials Lithium Hydride (LiH)
and Tungsten (W) for the neutron and gamma shielding, respectively.
Table 5.2.4.4-1 Shielding material properties
Lithium Hydride (LiH)
Density 0.78 g/cm3
Tungsten (W)
Density 19.3 g/cm3
As shown in the table above, we can see that the density of tungsten is very high and a
driving design factor for the shielding since too much tungsten shielding could lead to a very
high and undesirable mass.
A design criterion for the shielding is that radiation does not interact with the TPV power
conversion system since it will decrease the efficiency of the power conversion. Therefore, the
geometry of the shielding is driven by the diameter of the TPV and armor diameter of 4.4 m.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 130
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Geometry of the shielding considered are the simple cylindrical shape and the conical shape.
The cylindrical shaped shielding added unnecessary shielding material which increases
considerable amount of mass due to the high density material Tungsten. Therefore, we use the
conical shape for the shield for the considerable mass saving. Another advantage of the conical
shaped shielding is that the radiation will propagate outward away from the spacecraft in an
angle [9].
Also neutron exposure less than 0.1 mrem/hr is a desired characteristic [8] and it is a
conservative and safe value considering STVs do not have any human interaction until the crew
arrives to Ceres and in vicinity of the STV.
W-LiH-W configuration of layers is desirable for extra protection against strong gamma
radiation and a very safe against extra gamma radiation that surpasses the first line of shielding.
First layer of the Tungsten will dissipate most of the gamma radiation and second layer
consisting of Lithium Hydride will dissipate most of the neutron radiation. The third layer of
tungsten is added to shield the spacecraft from any neutron and especially gamma radiation that
passes the first two layers of shielding.
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Figure 5.2.4.4-1 Shielding geometry and layer configuration
Table 5.2.4.4-2 Shielding specifications
Materials / Configuration W-LiH-W
Lithium Hydride Thickness 20.5 cm
Tungsten Thickness 4 cm
Radiation Exposure <0.1 mrem/hr
Propagation Angle 19.837 degrees
Total Mass 18008.71 kg
Radiation
Radiation
Propagatio
n
Radiation
Propagatio
n
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Author: Jared Dietrich
5.2.4.5 Low Thrust Engine Power
The magnetoplasmadynamic thrusters (MPD) aboard each Supply Transfer Vehicle (STV)
produce 25 N of thrust. We design the power system requirement according to the amount of
total thrust supplied by MPDs. Figure 5.2.4.5-1 shows a plot of the power required for various
thrust and specific impulse combinations.
Figure 5.2.4.5-1 Power required for MPDTs per STV
We choose a specific impulse of 5000 seconds and an efficiency of 50% [1] for an MPDT.
Therefore, at a thrust of 25 N, we require 1.226 MW of power. The power is calculated in Eq.
5.2.4.1-1.
(5.2.4.5-1)
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where P is the power in kilowatts, F is the thrust in newtons, go is the acceleration due to gravity,
Isp is the specific impulse, and η is the efficiency [15]. The total power required for both STV 1
and STV 2 is 2.45 MW.
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References
[1] Ian M. Rousseau, Analysis of a High Temperature Supercritical Brayton Cycle for Space
Exploration, Research Science Institute, July 31, 2007.
[2] Mukund R. Patel, Spacecraft Power Systems, CRC Press, 2005.
[3] Lee S. Mason, Carlos D. Rodriguez, Barbara I. McKissock, SP-100 Reactor with Brayton
Conversion for Lunar Surface Applications, Ninth Symposium on Space Nuclear Power
Systems Albuquerque, New Mexico, January 12-16, 1992.
[4] Lee S. Mason, Power Technology Options for Nuclear Electric Propulsion, NASA Glenn
Research Center, 37th
Intersociety Energy Conversion Engineering Conference, 2002.
[5] Aaron E. Craft and Jeffrey C. King, Reactivity Control Schemes for Fast Spectrum Space
Nuclear Reactors, Missouri University of Science and Technology.
[7] Frank H. Welch, Lithium Hydride: A Space Age Shielding Material, Rockwell
International Corporation, Canoga Park, California, May, 1973.
[8] Anthony J. Hanford, Ph.D., Editor, Advanced Life Support Baseline Values and
Assumptions Document, Lockheed Martin Space Operations, 2004.
[9] L.W. Lee, Jr., Shielding Analysis of a Small Compact Space Nuclear Reactor, Air Force
Weapons Laboratory, August, 1987.
[10] Jasbir Singh, “Heat Transfer Fluids and Systems for Process and Energy Applications”,
Marcel Dekker, INC., New Yourk and Basel, 1985
[11] Mauk, M.G. and V.M. Andreev, GaSb-related Material for TPV Cells, Semiconductor
Science and Technology, 18 (2003), S191-S201.
[12] Gabler, H., Thermophotovoltaic Generation of Electricity, Proceeding of EuroSun’96,
1996.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 135
Author: Alex Park
[13] Baldasaro, P.F., Thermodynamic Analysis of Thermophotovoltaic Efficiency and Power
Density Tradeoffs, Journal of Applied Physics, Vol. 89, No. 6, 2001.
[14] Jasbir Singh, “Heat Transfer Fluids and Systems for Process and Energy Applications”,
Marcel Dekker, INC., New Yourk and Basel, 1985
[15] Martinez-Sanchez, M. and Pollard, J.E., “Spacecraft Electric Propulsion – An Overview,”
Journal of Propulsion and Power, Vol. 14, No. 5, Sep-Oct, 1998
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 136
Author: Jared Dietrich
5.2.5 Propulsion Systems Overview
The propulsion systems we design for both Supply Transfer Vehicles (STVs) consist of
chemical kick motors, chemical hover motors, magnetoplasmadynamic thrusters (MPD), and
attitude and control thrusters. The latter will be covered in section 5.2.6.
We employ chemical kick motors for earth departure from low earth orbit (LEO). Each STV
requires three kick motors. We design the motors based on parameters entered into Rocket
Propulsion Analysis [1]. Following the kick from LEO, the STVs employ a cluster of MPDs in
order to obtain a transfer orbit to Ceres. As the STVs reach Ceres, we fire a second kick from
the Ceres Regime Motors (CRM). The CRM kick places both STVs into a low Ceres orbit
(LCO). Finally, the STVs employ their respective CRM for a 60 second hover and landing on
the surface.
5.2.5.1 Earth Departure Kick Motor
We assume a chamber pressure 6 MPa for each kick motor. The propellants are liquid
hydrogen (LH2) and liquid oxygen (LOX). Analysis of the bipropellant assumes an optimum
oxidizer to fuel ratio of 5.136. We then determine the size and dimensions of the motor chamber
and nozzle. The length of the motor is 4.15 m and has a diameter of 2.14 m. Following a
systematic approach [2], results of the kick motor parameters are found in Table 5.2.5.1-1.
Table 5.2.5.1-1 Earth departure kick motor data
Vehicle Number of Motors Mass, kg Volume, m3
Thrust, kN Isp, sec
STV 1 3 7,967 73.2 1,500 458.3
STV 2 3 7,967 73.2 1,500 458.3
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The kick motors are mounted to each STV by an adapter. The adapters are 5 m long by 10 m
in diameter and have a mass of 3026 kg each [3]. After all propellants are consumed, the kick
motors and adapter are jettisoned from the STVs. Figure 5.2.5.1-1 shows a rendering of the
adapter attached to the bottom of STV 1. The adapter is transparent in order to illustrate the
landing legs, Ceres Regime Engine, and Kick Motors contained within.
Figure 5.2.5.1-1 STV kick motors attached to adapter
5.2.5.2 Magnetoplasmadynamic Thrusters (MPDs)
In this section, we will describe the process for selection of our low thrust engines. Once a
required thrust was supplied by the mission design group, power requirements and performance
characteristics were provided. We supply each STV with three MPDs. The propellant is liquid
hydrogen (LH2) and is heated until vaporization by the nuclear reactor. Performance
characteristics and dimensions of the MPDTs are found in Table 5.2.5.1-2.
By: Jared Dietrich
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 138
Author: Jared Dietrich
Table 5.2.5.1-2 Earth departure kick motor data
Vehicle Number of MPDTs Mass, kg Volume, m3
Thrust, kN Isp, sec
STV 1 3 1,992.0 0.005 25 5,000
STV 2 3 1,992.0 0.005 25 5,000
5.2.5.4 Ceres Regime Motors (CRMs)
In this section, we describe the design process for our hover and landing motor. We account
for Ceres’s gravitational acceleration when calculating the thrust requirements. We assume
negligible power requirements for both the kick motors and CRMs. A negligible power
assumption is valid since the MPDs, CRMs, and kick motors never fire at the same time.
Therefore, there is always enough power being supplied by the nuclear reactor.
Performance characteristics of the CRMs are found in Table 5.2.5.1-3. Each STV requires
only one CRM. They are mounted on the bottom of the core STV canister.
Table 5.2.5.1-3 Ceres Regime Motor Data
Vehicle Number of Engines Mass, kg Volume, m3
Thrust, kN Isp, sec
STV 1 1 181.2 0.070 100 452
STV 2 1 181.2 0.070 100 452
We assume a chamber pressure of 8 MPa. The propellants are liquid hydrogen (LH2) and
liquid oxygen (LOX). Analysis of the bipropellant assumes an optimum oxidizer to fuel ratio of
5.064. We then determine the size and dimensions of the motor chamber and nozzle. The length
of the motor is 4.15 m and has a diameter of 2.14 m.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 139
Author: Jared Dietrich
Fig. 5.2.5.4-1 CRE nozzle geometry from RPA
By: Jared Dietrich
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 140
Author: Jared Dietrich
Co-Author: Evan Helmeid
5.2.5.5 Ceres Regime Motors - Hover
We determine the engine thrust level by analyzing an appropriate minimum thrust needed to
descend to the surface, hover, and land. In principle, we need a thrust-to-weight ratio (T:W) of
~2 to maintain sufficient control of the vehicle upon descent, keeping in mind that the engine
must achieve a T:W of 1 to hover, and a T:W of <1 to actually land on the surface. The engine is
throttleable down to 10% of its nominal thrust, which allows us to achieve the necessary range of
T:W ratios. Table 5.2.5.5-1 outlines the necessary criteria and then engine capabilities.
Table 5.2.5.5-1 Required engine characteristics and actual engine specifications
STV 1 STV 2
Required Achievable Required Achievable
Thrust – Nominal (kN) - - 100.0 - - 80.00
Thrust – Min (kN) - - 10.00 - - 8.000
Weight – Max (kN) 37.69 - - 29.02 - -
Weight – Min (kN) 34.36 - - 26.41 - -
T:W Optimal Descent >2.0 2.6~2.9 >2.0 3.4~3.8
T:W Land <0.7 0.29 <0.7 0.38
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 141
Author: Jared Dietrich
Co-Author: Evan Helmeid
References
[1] Ponomarenko, Alexander, “RPA: Tool for Liquid Propellant Rocket Engine Analysis C++
Implementation”
[2] Humble, Ronald W., and Henry, Gary N., and Larson, Wiley J., Space Propulsion Analysis
and Design, The McGraw Hill Companies, Inc., New York, 1995
[3] Bednarcyk, Breatt, Arnold, Steven, Hopkins, Dale, “Design of Fiber Reinforced Foam
Sandwich Panels for Large Ares V Structural Applications”, Glenn Research Center, OH,
2010
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 142
Author: Paul Frakes
5.2.6 Attitude Determination and Control Systems (ADCS)
We equip each STV with an inertial Motion Reference Unit and associated computer system,
which serves as the attitude determination system. The attitude control system consists of six
attitude control thrusters, each attached to Canfield joints (see Fig. 5.2.6-1).
Figure 5.2.6-1. Model of a Canfield joint with an attitude control thruster. The payload of
the distal plate can be maneuvered through 2π steradians; the central propellant feed tubes
are flexible.
The Canfield joint is a new technology currently under development by Professor Stephen
Canfield at Tennessee Technological University. The design was selected for use on the Orion
Crew Module of NASA’s Constellation program. The design currently stands at a Technology
Readiness Level of only approximately 4, but we assume that the design will be human-rated in
time for our mission.
Canfield joints enable us to reduce the number of required attitude engines to six (from the
traditionally required minimum of sixteen) since the joints can be gimbaled to point in any
By: Alex Roth and Paul Frakes
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 143
Author: Paul Frakes
direction within a hemisphere, i.e., the range of motion of the distal plate is 2π steradians. The
joints are also controlled by the attitude determination computer system. Each Canfield joint is
controlled by three motors (not shown in Fig. 5.2.6-1) and is a two degree-of-freedom system.
The motors surround the base plate and control the three sets of jointed beams, allowing the
distal plate to move.
Note that the attitude determination computer system also controls the Canfield joint which
controls the pointing of the STVs’ communication dishes. See Section 5.2.8 for details.
The STVs are three-axis stabilized. We place the attitude control thrusters at the top and
bottom of three of the six payload shrouds that surround the core of the STV, as shown in Fig.
5.2.6-2. In this way, we couple the thrusters, allowing for pure translation along any axis, pure
rotation along any axis, and any combination of the two. We also space the thrusters evenly
around the vehicles, allowing 120° between them. This symmetry allows for the most efficient
use of attitude control propellant. The engines produce 20 Newtons of thrust each, and we
employ hypergolic propellants monomethylhydrazine (MMH) and nitrogen tetroxide (NTO). We
select hypergolic propellants over cryogenics because the cryogenics tend to boil off over the
long period of time over which we need to store the propellants. Each engine has an Isp of 220
seconds.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 144
Author: Paul Frakes
Figure 5.2.6-2. STV indicating the location of the attitude thrusters at the bottom of three
of the six payload shrouds surrounding the core of the vehicle. The other three thrusters
(located at the top of the configuration) are not shown.
The attitude control system is required to point each STV according to the steering law given
in Eq. 5.2.6-1 below [1].
(5.2.6-1)
This steering law says that the thrust of the STV’s main low-thrust engines must be in the
direction of the vehicle’s velocity.
The attitude control system is also required to correct for the effect of environmental torques
and forces that act on the spacecraft. These include solar radiation, solar wind, collisions with
particles in the Van Allen belt, atmospheric drag (during mission phases when this analysis is
appropriate), gravity gradient, and a variety of other effects [2]. We show that these “other
By: Jared Dietrich and Paul Frakes
Attitude Control Thrusters
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 145
Author: Paul Frakes
effects” (such as the effect of Earth’s magnetic field on each spacecraft as a function of the net
charge of the spacecraft) are negligible in a preliminary analysis, shown in the Appendix A.5.2.6.
Those effects that are specifically enumerated above, however, are taken into account in the full
analysis presented in Appendix A.5.2.6.
Finally, we must take into account the variations in spacecraft geometry and mass
distribution, namely as the center of mass and moments of inertia change throughout the constant
thrust phase of the mission. We account for all of these effects and provide appropriate
propellant masses to counter the effects.
The total mass, power, and volume requirements of STV1 and STV2 ADCS are given below
in Table 5.2.6-1. The following numbers take into account all of the attitude concerns for the
interplanetary transfer phase of the mission.
Table 5.2.6-1. Mass, power, and volume requirements of STV1 and STV2 ADCS
Mass, kg Power, kW Volume, m3
STV1 2,336.76 0.08 2.71
STV2 2,567.98 0.08 2.38
Total 4904.74 0.16 5.09
We provide a more detailed look at these numbers in Appendix A.5.2.6.
It is also important to note that we require additional attitude propellant throughout the STV
construction process in LEO, because atmospheric drag will cause the orbit of the STVs to decay
as we construct them. An additional 278.46 kg of propellant is required to keep STV1 in orbit
during construction, and 148.04 kg of propellant is required to keep STV2 in orbit during
construction. Even though STV1 is more massive than STV2, we do not begin construction of
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 146
Author: Paul Frakes
STV2 until several months after we begin construction of STV1, which explains why less total
propellant is needed to keep STV2 in LEO.
We also provide a small amount of control propellant for the landing phase, when the STVs
descend to the surface of Ceres from LEO. See Appendix A.5.2.6 for details.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 147
Author: Paul Frakes
References
[1] Sweetser, T. H., Cherng, M. J., Penzo, P. A., and Finlayson, P. A. “Watch Out, It’s Hot!
Earth Capture and Escape Spirals Using Solar Electric Propulsion,” AAS 01-439, 2001.
[2] Longuski, J. M., Todd, R. E., and König, W. W. “Survey of Nongravitational Forces and
Space Environmental Torques: Applied to the Galileo,” Journal of Guidance, Control, and
Dynamics, Vol. 15, No. 3, 1992.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 148
Author: Andrew Curtiss
5.2.7 Structures and Thermal Systems
5.2.7.1 Propellant Tanks
The propellant tanks hold the LH2 and LOX required for the stages of the mission. This
requirement is made up of kick propellant, Ceres orbit propellant, Ceres landing propellant, and
low thrust propellant. The kick, Ceres orbit, and Ceres landing propellants burn at a ratio of
5.136 parts LOX to 1 part LH2 while the low thrust propellant consists only of LH2. We use
these requirements, along with the dimensions of the Ares V payload bay, to define the required
size of the propellant tanks. An additional requirement on the propellant tanks is that they must
be large enough to satisfy the ISPP station propellant storage tank requirements. This measure
avoids the need to launch empty tanks for use at the ISPP stations.
Table 5.2.7.1-1 contains the dimensions of each of the propellant tanks which correspond to
Fig. 5.2.7.1-1. Both STVs use the same basic tank design and dimensions although they contain
different amounts of propellant. Because the tanks need to hold enough propellant to satisfy the
ISPP station requirements they are designed to be slightly larger than the size necessary to hold
the required propellant.
Table 5.2.7.1-1 Propellant tank parameters
Oxygen Tank
Radius, m Length, m Thickness, m Mass, kg Volume, m3 # of Tanks
3.5 0 0.0125 952.3979 181.2937 5
Hydrogen Tank
Radius, m Length, m Thickness, m Mass, kg Volume, m3 # of Tanks
4.27 14.5 0.0125 2249.7 1164.7 2
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Author: Andrew Curtiss
Figure 5.2.7.1-1 We model the propellant tanks as spheres and cylinders with
hemispherical ends, as shown here.
Table 5.2.7.1-1 lists the L and r dimensions shown in Fig. 5.2.7.1-1. The LOX tank has a
length of zero meters because it is a spherical tank as shown. The tanks are 12.5 mm thick and
are made up of carbon fiber and multilayer insulation. We design the thickness of the tank based
on the strength required to contain the 250,000 Pa of pressure inside the tank and also to insulate
the cryogenic propellant within.
Andrew Curtiss
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 150
Author: Andrew Curtiss
5.2.7.2 Landing Legs
We land the STV on the surface of Ceres with a set of four landing legs which are made
mostly out of carbon fiber. We model the legs with shock absorbers to help provide a soft
landing and to avoid excess stresses on the vehicle. These shock absorbers consist of a
compressible spring with the capability of absorbing the shock of a landing at 10 m/s. The shock
absorber consists of a strengthened steel spring which compresses to absorb landing energy. The
legs also feature landing pads which provide a stable foundation for the STV on uneven turf.
Figure 5.2.7.2-1 contains a dimensioned diagram of the assembled landing leg components.
There are four landing legs on each of the STVs. We position the legs around the bottom of
the center module at an angle of 30 degrees relative to the horizontal. We space the legs evenly
around the circumference of the center module. We design the legs so that at maximum load,
which occurs when the fully loaded STV hits the surface of Ceres at 10 m/s, the springs
compress halfway. An additional constraint on the design of the landing legs is that at maximum
compression there must be a clearance between the Ceres regime engine and the surface of
Ceres. The characteristics of the spring are listed in Table 5.2.7.2-1.
Table 5.2.7.2-1 Landing leg spring parameters
Force, N Spring Constant, N/m Turns Coil Diameter, m Wire Diameter, m
16991 23582 10 0.0986 0.0147
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 151
Author: Andrew Curtiss
Figure 5.2.7.2-1 The dimensions of landing leg components when the spring is
uncompressed.
Table 5.2.7.2-2 contains a listing of the four components of the landing leg system and the
corresponding mass and volume of each of the components. The total volume listed is for only
one of the four legs on each of the STV’s.
Table 5.2.7.2-2 Components of the landing legs
Component Material Mass. kg Volume. m3
Upper Leg Carbon Fiber 18.0046 0.0212
Lower Leg Carbon Fiber 39.5608 0.0096
Spring Strengthened Steel 4.0497 0.0039
Footpad Carbon Fiber 7.3435 0.000516
Totals 68.9586 0.035216
Andrew Curtiss
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 152
Author: Andrew Curtiss
5.2.7.3 Module Connectors
The module connectors are a critical component of the STV because they hold all of the
modules together and transfer propellant, coolant, and electrical power between the center
module, reactor, and propellant tanks. The modules consist of an electromagnet, coolant pipes,
and propellant transfer pipes. We wrap all of these components in a layer of Kevlar to hold the
connector together and also to protect the connector from damage from micrometeorites. Each
connector is 20 cm in length with a 1 cm thick Kevlar protective layer. Table 5.2.7.3-1 lists the
components of the module connectors and Fig. 5.2.7.3-1 shows a cross section of the module
connector.
Table 5.2.7.3-1 Component breakdown of module connectors
Part Component Material Mass, kg Power, w
1 Propellant Pipe Aluminum 10.0939 0
2 Coolant Pipe Copper 0.9331 0
3 Protection Layer Kevlar 7.238 0
4 Magnet Zinc 42.84 144
Figure 5.2.7.3-1 The cross section of module connector. Notice the electromagnet with
copper wire wrapped around the core, the two coolant pipes and the propellant transfer
pipe.
Andrew Curtiss
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 153
Author: Andrew Curtiss
5.2.7.4 Payload storage container
The payload storage container can be described as a structural shell that makes up the center
module of the STV. This shell contains all of the cargo items of the center module. The bottom
of the storage container features three kick engines and the Ceres regime engines. We place the
landing legs around the bottom of the container. The container fits in the upper portion of the
extended Ares V payload bay. The floor of the storage container supports the load of the payload
inside even during the high g loading experienced during launch. The material we chose for the
container is carbon fiber because of its relatively low density and high strength. Figure 5.2.7.4-1
features a dimensioned sketch of the storage container.
Figure 5.2.7.4-1 The dimensions and shape of the payload storage container.
Andrew Curtiss
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 154
Author: Andrew Curtiss
5.2.7.5 Food storage and radiation shielding containers
The food supply on the STV is vulnerable to radiation during the transfer flight to Ceres.
Therefore, we require radiation shielding container to protect the food supply. We arrange the
food in a cylindrical shape container covered with 20 g/cm2
of radiation protection. Figure
5.2.7.5-1 below shows a dimensioned picture of the food storage container that is on each STV.
It holds enough food to last 6 astronauts 3.5 years.
Table 5.2.7.5-1 Radiation Protection
Mass Cylinder, kg Mass w/ Food, kg Volume, m3
Food Radiation 15747 23912 5.8324
Protection
Figure 5.2.7.5-1 Dimensioned diagram of the radiation protection food storage container.
Notice that the height is twice the radius. This constraint minimizes the surface area of the
container.
Andrew Curtiss
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 155
Author: Kim Madden
5.2.7.6 Thermal Control System
Table 5.2.7.6-1 shows a compiled chart of the mass, power, and volume requirements for the
STV thermal control system.
Table 5.2.7.6-1: STV Thermal Control System Summary
Component Mass, kg Power, kW Volume, m3
MLI Covering 139.50 0 0.50
Heat Pipe 169.11 0 0.98
Radiators 551.87 0.08 0.17
Aluminum Plates 21.08 0 0.008
Heater 179.03 0.08 0.18
TOTAL 1060.58 0.16 1.86
A detailed description of the thermal control system for the STV can be found in Appendix
A.5.7.6.2.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 156
Author: Sarah Jo DeFini
Co-author: Paul Frakes
5.2.8 Communication Systems
Our Supply Transfer Vehicle Tracking, Telemetry, and Command (TT&C) Communications
system is responsible for the health and status of 14 different subsystems, which totals to up to
180 different health and status signals. During nominal operation, each subsystem sends health
and status signals to the general processing computer once every hour. The computer formats
this data and, once a month, sends all of its stored data to the external communications hardware.
At this time, the telemetry dish aligns itself so that it can transmit data to a visible a Tracking and
Data Relay Satellite, which relays data to the Deep Space Network on Earth. Earth can then
transmit data to each vehicle to correct for any status irregularities. We limit total monthly
contact time to one hour.
5.2.8.2 Antenna Pointing Accuracy
The communication dish must be pointed within 1 degree of its target during all transmission
times. This presents a challenge for a dish that is statically anchored to the STV, since the
vehicle must be pointed according to the vehicle’s steering law, which in general does not allow
for correct communication dish pointing. This problem can be avoided, however, by using a
Canfield joint, depicted in Fig. 5.2.8.2-1.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 157
Author: Sarah Jo DeFini
Co-author: Paul Frakes
Figure 5.2.8.2-1. Model of the Canfield joint, depicted here with a thruster instead of a
dish. The communication dish is mounted to the distal plate, which allows accurate dish
pointing.
The communication dish is mounted to the distal plate of the Canfield joint, which enables
pointing through 2π steradians (i.e., hemispherical range of motion). Since the STVs are each
three-axis stabilized (i.e., no spin is required), the dish can be placed on the side of the vehicle
that will face towards the Earth the whole time, alleviating any concern that communication will
be impossible because of vehicle roll.
The Canfield joint will be controlled by the same computer system that controls attitude, since
the attitude thrusters are also controlled by Canfield joints. See Section 5.2.6 for a detailed
description of that computer system.
By: Alex Roth and Paul Frakes
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 158
Author: Sarah Jo DeFini
5.2.8.3 Space Communications and Navigation (SCaN)
Space Communications and Navigation (SCaN) is NASA’s program to coordinate the Deep
Space Network (DSN) and the Tracking and Data Relay Satellite system (TDRSs) with each
other and with other low-Earth and ground networks [1]. Using these networks makes it possible
for each STV to communicate with Earth at any point during the transit from Earth to Ceres.
Once a month, each STV aligns itself to communicate with one of NASA’s Tracking and Data
Relay Satellites, which have dish diameters of at least 15 meters and transmit data to the Deep
Space Network on Earth’s surface. For more information about SCaN see Appendix 5.2.8.3.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 159
Author: Sarah Jo DeFini
5.2.8.4. Link Budget
We use the S-band for both our downlink (2.4 GHz) and uplink (2.85 GHz) frequencies. This
is standard for tracking, telemetry, and command data. We also needed direct communication
with earth to be possible even at the apoapsis of the transfer orbit, so we set the propagation path
length equal to the maximum distance from Ceres to Earth during STV transfer. The remaining
link budget input parameters are summarized in Table 5.2.8.5-1. The link budgets for the uplink
and downlink were calculated using the same process outlined in Appendix D.1.1.5. A more
complete link budget is included in Appendix A.5.2.8.4.
Table 5.2.8.4-1: STV Link Budget Parameters
Downlink Uplink
Frequency (GHz) 2.4 2.85
Data Rate (bps) 5000 5000
Input Power (kW) 8.5 8.5
Pointing Loss (degrees) 0.021 0.021
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 160
Author: Sarah Jo DeFini
5.2.8.5 Communication Hardware
Each system sends its health and status signals to a general processing computer, which stores
all of this information. The computer we chose is similar to one of the radiation-hardened
computers currently in use on the Space Shuttle. Its parameters are summarized in Table 5.2.8.6-
1.
Table 5.2.8.5-1: General Processing Computer Specifications
Parameter Value Units
Mass 29 kg
Power 550 W
Volume 0.025 m3
The communication dish was also sized using the previously described process. Its physical
parameters are summarized in Table 5.2.8.6-2. Trade-off studies are included in Appendix
5.2.8.5.
Table 5.2.8.5-2: Communications Dish Physical Parameters
STV Communications Dish
Diameter (m) 1
Mass (kg) 18.63
Power (kW) 8.5
Volume (m3) 0.005
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 161
Author: Sarah Jo DeFini
References
[1] “NASA’s Mission Operations and Communications Services,” [online database],
http://deepspace.jpl.nasa.gov/advmiss/docs/NASA_MO&CS.pdf [retrieved 1 February 2011].
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 162
Author: Evan Helmeid
5.2.9 Ceres Operations
5.2.9.1 Landing on Ceres
To land the two Supply Transfer Vehicles (STVs) on the surface, we propagate an optimal
trajectory using a two-point boundary value problem solver in MATLAB (see appendix F.4.2 for
details). The trajectory minimizes time using x- and y-position and velocity as process equations,
a flat-surface assumption, and a constant mass assumption. We show the resultant trajectory in
Fig. 5.2.9.1-1. The details of this analysis are discussed in section A.5.2.9.1.
Figure 5.2.9.1-1 Optimal landing trajectory for STV1 (left) and STV2 (right) to descend
from LCO to the equator of Ceres using a flat-surface model, constant mass, and constant
thrust. The trajectory minimizes time.
In the final trajectory, note that the steering angle ends with the spacecraft at near vertical,
which is necessary for a feasible landing.
To account for unexpected surface features, such as boulders, and to allow the STV an
amount of buffer in landing in a desired location, we add enough propellant to allow the
0 2 4 6 8 10 12
x 104
0
1
2
3
4
5x 10
4
Evan Helmeid
Trajectory of Spacecraft
X-position (m)
Y-p
ositio
n (
m)
Final altitude
Final trajectory
0 100 200 300 400 500 600-100
-50
0
50
100
Evan Helmeid
Steering Angle, , vs Time
time (s)
Ste
ering A
ngle
,
(deg)
50 100 150 200 250 300 350-150
-100
-50
0
Evan Helmeid
Y-Velocity vs X-Velocity
Velocity x-component (m/s)
Velo
city y
-com
ponent
(m/s
)
0 100 200 300 400 500 600-200
-100
0
100
200
300
400
Evan Helmeid
Velocity Components vs Time
time (s)
Velo
city (
m/s
)
Velocity: x-direction
Velocity: y-direction
0 2 4 6 8 10 12
x 104
0
1
2
3
4
5x 10
4
Evan Helmeid
Trajectory of Spacecraft
X-position (m)
Y-p
ositio
n (
m)
Final altitude
Final trajectory
0 100 200 300 400 500 600-100
-50
0
50
100
Evan Helmeid
Steering Angle, , vs Time
time (s)
Ste
ering A
ngle
,
(deg)
50 100 150 200 250 300 350-200
-150
-100
-50
0
50
Evan Helmeid
Y-Velocity vs X-Velocity
Velocity x-component (m/s)
Velo
city y
-com
ponent
(m/s
)
0 100 200 300 400 500 600-200
-100
0
100
200
300
400
Evan Helmeid
Velocity Components vs Time
time (s)
Velo
city (
m/s
)
Velocity: x-direction
Velocity: y-direction
By: Evan Helmeid
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 163
Author: Evan Helmeid
spacecraft to hover for 60 seconds before landing. Final landing parameters are presented in
Table 5.2.9.1-1.
Table 5.2.9.1-1 Summary of STV1 and STV2 landings, including propellant requirements,
engine thrust levels, and trajectory characteristics
STV 1 STV 2 Units
Wet mass (in LCO) 140.1 107.9 T
Dry mass 125.3 95.98 T
Propellant mass 12.82 10.04 T
Thrust range 100~1000 80~800 kN
Tnominal:W range 2.6~2.9 3.4~3.8 - -
Burn time 578.9+60 562.7+60 s
The optimal landing of both STVs on the surface of Ceres requires minimal propellant due to
the use of an appropriately-sized engine and an optimal trajectory, but maintains a factor of
safety because of its hover capabilities and thrust buffer. The factors taken into account when
sizing lend to an overall system that meets requirements and achieves low mass due to
appropriate, interrelated sizing.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 164
Author: Andrew Curtiss
5.2.9.2. Deployment of Cargo on Ceres
Upon arrival on the surface of Ceres, the various cargo components in the modules must be
deployed. We deploy the communication satellites when the STV reaches LCO and the rest of
the vehicle lands without the satellites. After the STV has landed on the surface of Ceres, the
other cargo deploys including the exploration and rescue rovers, the ISPP station harvesters,
ovens, and reactors. To separate the modules and deploy the cargo we simply cut the power to
the electromagnets. This action separates the modules and allows the cargo to be deployed. For
simplicity, the propellant tanks remain attached to the center module which becomes the
propellant production station. The rescue rover flies out of the top of the STV while the
exploration rovers and the harvesters drive out of the STV.
Detailed Vehicle Descriptions Supply Transfer Vehicle Page 165
Author: Sonia Teran
5.2.10 End of Life Configuration
Our supply vehicles have a one-way journey. Once we land them on Ceres, they complete
their mission as transfer vehicles. They will not be reused for future transfer missions and,
therefore, their lives end on Ceres. We do reuse a majority of our components for the ISPP
stations. Our propellant tanks are reused as storage tanks for the ISPP stations. The center section
is also retained intact for the ISPP facility. Those components which are no longer used are the
storage containers once the supplies are removed; these containers will remain on Ceres. Also,
our reactor is no longer needed and will be throttled down and shut off. At this point the STVs
will be used mainly for ISPP requirements.
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5.3 Crew Launch Vehicle
As mentioned previously, two different types of vehicles deliver all payloads and our six-
member crew into low Earth orbit (LEO). The Ares V brings propellant, tanks, and the Crew
Transfer Vehicle (CTV) structure. An Ares I transports the crew safely to LEO in their Capsule,
as a fully constructed CTV awaits their arrival. The mass and volume breakdowns as well as the
design intent behind specific components of the CTV are discussed in the crew launch manifest.
5.3.1 Launching the Crew
Our crew arrives in LEO aboard the Crew Capsule by way of an Ares I. The Ares I launch
vehicle lifts only the mass of the Capsule and the mass of the six-member crew from Earth’s
surface. The only adaptation necessary for the crew launch vehicle is an extended spacecraft
adapter between the Crew Capsule and the upper stage. The spacecraft adapter is currently
designed for the diameter of the Orion capsule of 5.02 m [1], but our Capsule has a slightly
larger diameter of 5.25m. Despite this modification, the Ares 1 fits the payload capabilities that
our Crew Capsule needs. The Ares I has a payload mass capacity of 25 metric tons and is able to
lift 4120 m3 as payload [2]. Our Ares I carries 9.8 metric tons and 27.5 m
3 into LEO, well below
the limits of this launch vehicle.
In order to improve the probability of success for our mission, the crew launch vehicle
contains an important contingency which increases our crew’s safety. For this contingency plan
we use the launch abort system (LAS) designed by Orbital Science. This LAS allows the crew to
escape any harm from a malfunction with the lower stages during launch, by jettisoning away
from the Ares I. The only modification necessary for current designs of the LAS system is the
connection between our Crew Capsule and the LAS system. A difference in sizes of the upper
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portion of our Crew Capsule and the Orion capsule exists and connection radius of the LAS must
be modified. An image of Orbital’s LAS is found in Fig. 5.3.1-1.
Figure 5.3.1-1 Configuration of Orbital’s LAS that will be used for crew launch. Pad Abort
1 Test Configuration – Orbital [3].
5.3.2 Crew Launch Manifest and Timeline
The total injected mass to LEO (IMLEO) we launch is 1,073,978 kg, and the total injected
volume to LEO (IVLEO) for the CTV is 3,997 m3. Launching all components of the CTV not
only requires 2 different launch vehicles, but also 4 different kinds of launches. Each type of
launch contains a different payload volume and mass distribution. The first launch type, Type A,
will also be the first launch for the CTV and will consist of mainly the structural and power
components of the CTV. Launches B and C contain propellant tanks, engines, and propellant.
Launch D brings our crew and the Crew Capsule to LEO. The total IMLEO and IVLEO for each
payload type are summarized in Table 5.3.2-1.
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Table 5.3.2-1 Launch type with corresponding mass and volume per launch.
Launch Type Mass, kg Volume , m3
A 129,728 411.3
B 123,950 497.6
C 187,248 688.4
D 9,836 27.5
We pack the most variety of payload components in payload launch. The breakdown for
launch Type A payload components can be found in Table 5.3.2-2.
Table 5.3.2-2 Launch A payload components for CTV launch.
Launch A: Payload
Crew Cabin
Storage Attic
CTV Elevator Structure
Tether and Low Thrust Engines
Reactor for CTV
Ballute System
Attitude and Control Systems
Within each payload component of the Type A payload configuration there are many
subsystems. Launching the crew cabin includes launching the entire crew living quarters, as well
as all human factors systems, for example the hygiene system, waste system, and health care
system. Along with the crew cabin, we send the crew storage attic in this launch vehicle. The
storage attic contains thermal radiators, food, the hydroponics system and the water regeneration
system. More structural components included in this launch are the CTV elevator shaft and the
main tether with the low thrust engines. Only one Ares V launch is necessary to bring all Type A
payload to LEO.
Primary tanks that travel with us for the entirety of the trip to Ceres are launched in the Type
B payload configuration. This payload configuration consists of a primary LH2 tank filled with
propellant, a primary LOX tank filled with propellant, and a high thrust engine attached to the
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tanks. Type C launches our earth departure tanks, propellant and engines. A total of 3 Type B
and 3 Type C launches bring all propellant and engines to LEO aboard Ares V launch vehicles.
The final launch type is Type D. This payload configuration consists of the Crew Capsule and
our 6 member crew. One launch of an Ares I transports the crew and the Capsule to the CTV in
LEO.
The CTV launch timeline begins with launch A on 3/11/2027. Each consecutive launch
occurs every 2 months after launch A, adding to the entire structure of the CTV. Our crew will
launch on 5/11/2028 for the payload Type D and the final launch of the CTV payload. The last
possible launch date is 8/11/2028 before our mission timeline would have to change. Our current
detailed launch timeline allows for a 3 month buffer between or final scheduled launch and the
last possible launch date. The detailed timeline our launches for the CTV are located in Table
5.3.2-3.
Table 5.3.2-3 CTV detailed launch timeline
Launch Type Launch Date
Launch A 3/11/2027
Launch B1 5/11/2027
Launch B2 7/11/2027
Launch B3 9/11/2027
Launch C1 11/11/2027
Launch C2 1/11/2028
Launch C3 3/11/2028
Launch D 5/11/2028
With a total of 8 launches over a period of nearly 1.5 years, we are able to combine all
IMLEO and IVLEO to form a complete CTV. The final launch in the sequence of launches
brings our crew to LEO to rendezvous with the completed CTV. Using the large payload
capabilities provided to us with the Ares V, we are able to do this in a timely matter.
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References
[1] Hatfield, S., “Project Orion Overview and Prime Contractor Announcement”, URL:
http://www.nasa.gov/pdf/156298main_orion_handout.pdf [cited 25 March 2011].
[2] “Overview: Ares I Crew Launch Vehicle,” Constellation Program, cited 25March 2011.
[http://www.nasa.gov/mission_pages/constellation/ares/aresl/index.html]
[3] “Orion Crew Exploration Vehicle,” Launch Abort System (LAS) Fact Sheet, cited
25March 2011.
[http://www.orbital.com/NewsInfo/Publications/Las_Fact.pdf]
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5.4. Crew Transfer Vehicle
5.4.1. Construction in LEO
We require a series of eight launches to assemble the Crew Transfer Vehicle in LEO due to
the size and mass of the components. Seven Ares V rockets and an Ares I rocket carry out each
of the eight launches with their respective payloads. Table 5.4.1-1 shows a schedule of the
launches up to crew boarding and habitation.
Table 5.4.1-1 – Construction timeline for the Crew Transfer Vehicle
Launch Number Launch Vehicle Vehicle Section
1 Ares V Center Chassis Section
2 Ares V Primary Tank 1
3 Ares V Primary Tank 2
4 Ares V Primary Tank 3
5 Ares V Earth Departure Tank 1
6 Ares V Earth Departure Tank 2
7 Ares V Earth Departure Tank 3
8 Ares I Crew and Capsule
When it comes to vehicle and component design, we mind the importance of the construction
procedure. We simplify the assembly by attaching prefabricated components to a central
structure in low Earth orbit. The center chassis acts as the housing for the effective crew payload
and other inert components. This center chassis contains the Connection devices for the
removable propellant tanks and engines. After the launch of each section of propellant tanks and
their engines to orbit, autonomous rendezvous takes place and connects to the vehicle chassis.
We carry out this procedure one by one until installation of all six propellant tanks. Once the
CTV reaches working condition, the crew and crew capsule launch to rendezvous with the
vehicle. Figures 5.4.1-1 through 5.4.1-3 show the vehicle configurations after the successive
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construction stages. The vehicle consists of the center chassis, the connected primary tanks, and
the Earth departure tanks.
Figure 5.4.1-1 Center chassis structure of CTV
By: Alex Roth
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Figure 5.4.1-1 shows only the center section of the crew transfer vehicle. This contains the
living quarters, CTV heat shield, communication systems, power source, low thrust electric
motors, tether system, the FORSE, and the connecting structure with all propellant tank mounts.
Figure 5.4.1-2 Assembled vehicle with primary tanks
By: Alex Roth
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The following three launches deliver the primary propellant tanks to LEO and connect to the
existing main structure. The ―primaries‖ have liquid oxygen (LOX) and liquid hydrogen (LH2)
tanks, a single high thrust kick engine, and a single lower thrust Ceres regime engine.
Figure 5.4.1-3 Assembled vehicle with all propellant tanks
The remaining three Ares V launches deliver the Earth departure, consisting of LOX tanks,
LH2 tanks, and a single high thrust kick engine.
By: Alex Roth
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5.4.2. Configuration Overview
The multiple tasks required of the Crew Transfer Vehicle throughout the mission cause the
vehicle to undergo various configurations. From a single configuration, the vehicle completes a
set of tasks during a particular stage in the mission. Section 3.1 introduces each vehicle
configuration and includes compact configuration, initial docking configuration, and extended
configuration. Each configuration has variations that are specific to certain stages in the mission.
Multiple variations of the compact configuration occur throughout the mission and we list
them below:
1) Stowed Capsule
2) Stowed Earth Departure
3) Ceres operations
Figure 5.4.2-1 displays each configuration. For the stowed capsule configuration, we place
the capsule at the top of the center chassis section to secure it for maneuvers and transit. This
configuration never locates the crew in the capsule. We classify the stowed Earth departure
configuration the same as the stowed capsule configuration except that all six propellant tanks
attach to the vehicle in a circular array. After the Earth departure burn, we jettison three tanks
leaving the three primary propellant tanks attached. Without Earth departure tanks in place the
capsule docks directly with the side port on the crew habitat without any cable extension. We
require this configuration for maneuvers on and at Ceres. We classify this mode as the Ceres
operations configuration. The tether does not extend by any amount for all of the compact
configurations indicating the commonality between them.
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Figure 5.4.2-1 Stowed capsule, Earth departure, and Ceres operations compact
configurations
Initial docking defines the second configuration; we show this in Fig. 5.4.2-2. In this set up,
the vehicle has all six propellant tanks attached and the tethers extend enough to allow the crew
By: Alex Roth
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capsule to dock at the side port on the crew habitat. This configuration occurs when the crew
arrives in the capsule and only once throughout the mission.
Figure 5.4.2-2 Initial crew docking configuration upon arrival of eighth launch. Notice the
green capsule attached to the habitat.
We extended the tethers out to a large distance for the third and final configuration as shown
in Fig. 5.4.2-3. The tethers extend during the transfer between Earth and Ceres where we require
By: Alex Roth
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the vehicle to spin for artificial gravity. The lengths of the tethers change based on the mass
distributions of the vehicle, which vary during the interplanetary transit. In addition, the radiator
panels extend to reject heat from the reactor while operating at maximum power output.
Figure 5.4.2-3 The crew habitat extends a distance to allow for artificial gravity
By: Alex Roth
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5.4.3 Outbound Trajectory
The goals for the outbound trajectory of the CTV include the following:
Reach Ceres within the mission requirement of a 2 year time limit for the crew time in
transfer
Provide the trajectory with the lowest propellant cost
These goals are the backbone to our selection of the final outbound trajectory to Ceres.
Appendix Section A.5.4.3 provides information on the detailed process used to design and select
our final trajectory for the CTV.
In order to achieve the goals of the CTV trajectory, a few key assumptions are made. We
assume both Ceres’s and Earth’s orbits are circular and coplanar. This assumption allows for
simplified calculations for this preliminary feasibility study. Assuming circular orbits is a valid
assumption to make, seeing as the eccentricity of Ceres’s orbit is 0.079 and Earth’s orbit is
0.0164 [1]; a truly circular orbit has an eccentricity of zero. The coplanar assumption provides
simplicity in calculations. In actuality, Ceres’s orbit is at an inclination of 10° relative to Earth’s
orbit. For a more detailed analysis, this inclination change must be taken into account to provide
the most accurate trajectory. Another important assumption for our analysis is that all ∆V’s are
considered impulsive. The addition of burn arcs in the analysis would provide a more accurate
description of the feasibility of this mission.
For the outbound trajectory of the CTV, we decide to use a combination of two types of
engines to transport us to Ceres. A chemical engine produces impulsive ∆V’s, one at Earth and
another at Ceres. Low thrust engines, Magnetoplasmadynamic (MPD), provide constant thrust
throughout the remainder of the transfer to Ceres. This combination of engines allows us to reach
Ceres within the time requirement.
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Starting in a LEO orbit of 350 km, we perform an impulsive ∆V to escape Earth’s gravity
field. Examining different trajectory options for the spacecraft at departure, we are able to vary
the flight path angle to find a ∆V which reduces propellant cost for this portion of the mission. A
more detailed discussion on this can be found in the appendix section A.5.4.3. Since Earth’s orbit
is assumed to be circular, we can calculate a ∆V by using Eq. 5.4.3-1. Where V∞ is our excess
velocity relative to Earth, rLEO is the radius from the center of the Earth to LEO, and µEarth is the
gravitational constant of Earth.
(5.4.3-1)
With a V∞ of 6.05 km/s we determine our ∆V at Earth to be 4.75 km/s. We find the mass of
propellant used during this ∆V by rearranging the ideal rocket equation. The Isp for the chemical
engine is 458 seconds and the gravitational constant, g0, is 9.805655 m/s2. The initial mass for
this calculation, m0, is the mass of the CTV in LEO including all propellant. Rearranging the
ideal rocket equation, we find that the mass of propellant for a specific burn is determined by Eq.
5.4.3-2.
(5.4.3-2)
Having such a large ∆V at Earth corresponds to a large amount of propellant used for that
burn. In order to reduce the overall mass of the CTV as it transfers, we jettison the tanks that
carry the propellant used for the ∆V at Earth departure.
Once the Earth departure tanks are jettisoned from the CTV, we begin our low thrust transfer
to Ceres. All 4 MPD thrusters are turned on and produce 33 N of thrust. We then thrust in the
direction of our instantaneous velocity vector throughout the transfer to Ceres. The state
equations of motion we use to numerically propagate this portion of the trajectory can be seen in
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Eq. 5.4.3-3 through Eq. 5.4.6-7. Where T is thrust, m is the mass of the CTV, Vr is the velocity of
the CTV in the radial direction, Vθ is the velocity of the CTV in the tangential direction, and r is
the position of the CTV with respect to the Sun. A more detailed discussion of the equations of
motion as well as the coordinate system we use can be found in A.5.4.3.
(5.4.3-3)
(5.4.3-4)
(5.4.3-5)
(5.4.3-6)
(5.4.3-7)
Since the MPDs are continuously thrusting throughout the transfer, we are constantly burning
propellant. The change in mass over time for the CTV can be seen in Eq. 5.4.3-8 where is the
mass flow rate of the engines, is the initial mass of the CTV in the heliocentric transfer and
dt is the change in time.
(5.4.3-8)
After 1.4 years of thrusting, we arrive at Ceres’s orbit with a V∞ of 2.48 km/s. Once at Ceres,
we perform a ∆V of 2.19 km/s to capture the CTV into a parking orbit altitude of 50 km above
the surface of Ceres. We calculate this value using Eq. 5.4.3-9 where instead of Earth conditions
we now use µCeres, Ceres’s gravitational constant, and rLCO the radius of LCO from the center of
Ceres. We also determine mass of propellant used during the burn with the previously mentioned
rocket equation, Eq. 5.4.3-2.
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(5.4.3-12)
A representation of the outbound transfer to Ceres can be found in Fig. 5.4.3-1. The red line
shows the low thrust portion of the mission. The impulsive occur at each celestial body.
Figure 5.4.3.-1 The CTV transfer from Earth to Ceres
The propellant costs throughout the outbound trip of CTV can be found in Table 5.4.3-1. To
take into account an increase in propellant used during a burn arc, as opposed to a purely
impulsive burn for our assumption, we increase each calculated mass of propellant for each ∆V
by 15%.
By: Trieste Signorino
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Table 5.4.3-1 Propellant masses used for each phase of the CTV outbound trip
Mission Phase Mass of Propellant, kg
∆V for Earth Departure 728,716
Low Thrust Transfer 28,941
∆V for Ceres Capture 134,700
We have now positioned the CTV at Ceres so it can descend to the surface and begin all
operations to fulfill the rest of the mission requirements. All of the goals for the outbound
trajectory of the CTV were accomplished. We arrive at Ceres and reduce the time of flight for
the transfer below the 2 year requirement to 1.4 years. This trajectory provides a non-optimal
solution to the problem, but was chosen to provide the minimum amount of propellant cost after
examining a select number of cases. A more detailed analysis including burn arcs, non-circular
and coplanar orbits, along with using optimization techniques, would produce the best solution
for this transfer. The large quantities of propellant mass could be reduced by replacing the
chemical engines with nuclear thermal engines. A comparison between nuclear thermal engines
and the chemical engines we use can be found in Appendix H.
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References
[1] ―HORIZONS Web Interface,‖ Solar System Dynamics, URL:
http://ssd.jpl.nasa.gov/horizons.cgi#results [cited 26 March 2011].
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5.4.4 Power Systems
Scope
An efficient, cost effective, and mass effective power system is a crucial component to the
mission’s success; without it nothing can function. In order to find the best means of meeting the
vehicles specific power requirements, we conducted several trade studies that took into account
everything regarding different types of power sources along with every detail that each option
brings to the table. In the end, the option that was chosen exhibited the lowest specific power in
terms of kg/kW and therefore the lowest cost in terms of cost per kilogram.
Background
At the conclusion of our analysis, we decided to use a nuclear fission reactor as the power
source for the Crew Transfer Vehicle (CTV). Using reactors for space applications is not a new
technology by any means. There have been multiple missions that incorporated these sorts of
power systems; for example the Jupiter Icy Moons Orbiter (JIMO) missions, along with Projects
Prometheus, Orion and Daedalus all experimented with nuclear powered spacecraft. The model
for our reactor is based similarly to the most recent U.S designed reactor, the SP-100 based on its
performance and reliability.
Risks
Anytime the topic of nuclear power comes up, there is almost just as much concern over the
potential dangers as there is the potential high performance capabilities. This is because nuclear
reactors produce very significant, and potentially deadly, amounts of radiation in the process of
producing useful electrical power. However, after extensive analysis and design options
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regarding the shielding around the core, we can safely say the crew will not be harmed by the
potentially dangerous levels of radiation produced.
Design/Analysis
Onboard the CTV, everything from the food storage systems and hydroponics to the low
thrust MPD thrusters requires a certain amount of power. All of the individual power
requirements can be seen below in Table 5.4.4-1:
Table 5.4.4-1 CTV power budget
Component Power Requirement, kWe
Food System 17
Recreation 2
House Cleaning 1
Maintenance System 2
Health Care System 2
Personal Communication Devices 1
Air Filtration/Recycling System 16
Air Circulation/Ducting 2
Communication Dish/System 11
Freezer(s) 2
Hydroponics 2
Water Regeneration System 0.23
Alternate Control Devices (CMG’s) 0.6
MPD Thrusters 1960
Total 2020
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After deciding on an official total power requirement of a little over two mega-watts, we
concluded that a 2.1MW nuclear fission reactor would be the best option for the application.
Selecting a nuclear reactor as the power system is advantageous for several reasons. For one,
this fission reactor has a much lower specific power (12.86 kg/kWe) than any other power
system option that is capable of producing such a significant power output, therefore making it
much more cost effective when it comes to cost per kilogram. And when it comes to packaging,
the total volume is about eight cubic meters which comfortably fits inside the Ares V shroud.
Nuclear Reactor Design Specifics
To satisfy all of the CTV’s power requirements, we use a single 2.1MW Uranium Oxide
(UO2) fueled, sodium-potassium pumped cooled nuclear reactor with dual Stirling power
converters and pumped water heat rejection system [2, 3]. The reactor is capable of running at
full power for up to eight years but will be throttled back, by way of Beryllium Oxide (BeO)
axial neutron reflectors, to roughly 100kWe once we land on the surface of Ceres [4]. This
reduction is because the MPDs will not be in use. The reactor core itself is made up of 1800
UO2 fuel pins and operates at an average temperature of roughly 955 K. The coolant, liquid
Sodium Potassium (NaK), enters the core at roughly 840 K and exits the core, entering the Heat-
Pipe System (HPS) at about 890 K [2]. We chose this particular coolant due largely because of
its extensive use in previous applications and its low freezing point (262 K), therefore requiring
little to no heating in space. The HPS is a type of heat rejection system that transports the heated
working fluid thru pipes inside large radiator panels that then absorb the heat and radiate it out to
space. For this specific reactor we chose a carbon-carbon composite as our panel material based
on its thermal and structural characteristics. In order to calculate the panel area needed to
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dissipate all of the excess thermal heat that the reactor generates, Eq. 5.4.4-1 below is
manipulated to output the required area.
(5.4.4-1)
is the thermal output we need to radiate which can be conservatively assumed to be
four times the usable electrical power generated (25% efficient). is the expected solar flux
(1400 W), is the emissivity of the carbon-carbon (.85), is the Stefan-Boltzmann constant
(5.6704e-8
), Tliquid is the temperature at which the coolant exits the core (890 K),
and Tatm is the surface temperature of Ceres (167 K). After inputting the given values into the
equation, the required area to dissipate 8.4MW of thermal energy is roughly 280 m2. This value,
along with the inputted values and reactor dimensions are tabulated in Table A.5.4.4-1 and Table
A.5.4.4-2 in appendix A.5.4.4. These panels will initially fold up on the sides of the reactor
vessel until the low thrust motors are activated. Only then will they unfold and extend outward
axially. Figure below illustrates just how the panels will look while folded up and also unfolded:
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Figure 5.4.4-1 Heat rejection panels in both the stowed and deployed orientations
Radiation Shield
There are several things we considered when we designed a space reactor radiation shield,
including: the magnitude of the radiation source, the types of radiation produced, the
configuration of the source, the payload, and dose limit. In our case, the fission reactor produces
both neutron and gamma radiation, meaning we need at least two different shielding materials.
The geometry of our reactor is cylindrical, so the shielding will simply be oriented concentrically
around its sides. The sensitive payload is a crew of six astronauts, and their dose limit, as
specified by the National Council on Radiation Protection is 50 rem/yr [1]. However, to err on
the side of caution, we designed a shield that will emit roughly 5-10 rems/yr to the astronauts. As
mentioned earlier, the reactor will produce both gamma and neutron radiation that must be
shielded. Neutrons are typically shielded by first reducing the energy through scattering, then
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absorbing the thermal neutrons. Materials such as hydrogen are very effective for scattering
neutrons, and for this reason, materials with high concentrations of hydrogen such as the material
we use, lithium hydride, are desirable for neutron shielding. Gamma radiation, on the other hand,
is attenuated by interactions with electrons through the photoelectric effect. Therefore, materials
with high electron densities are desirable for gamma shielding. When it comes to picking a high
density material, it comes down to a trade-off between effectiveness and mass allocation.
Obviously, the higher the density the material is, the higher the mass and the effectiveness is, so
we believe that tungsten is the best choice. Figure 5.4.4-1 below shows the reactor along with the
designed radiation shielding:
Figure 5.4.4-1 Reactor core (red) surrounded by radiation shielding – LiH (brown) and W
(blue)
1.79 m
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In the figure above, we can see just how thick the shielding is relative to the reactor core
itself. Table A.5.4.4-1 in appendix A.5.4.4 tabulates the material thicknesses for each layer of
shielding. Deciding of the location of the gamma attenuating material, tungsten was a trade-off
study in its own right. Being the denser of the two materials, it would make sense mass-wise to
position the tungsten just around the core. However, the fast moving neutrons coming from the
core create secondary gammas within the tungsten layer through inelastic scattering. To
minimize the dose contribution from these secondary gammas, the fast neutrons produced by the
reactor need to be thermalized through scattering by a highly hydrogen concentrated material
like lithium hydride before encountering the tungsten layer. This implies that placing the neutron
shielding materials between the core and the tungsten layer would reduce the secondary gamma
production produced in the tungsten. As expected, the secondary gamma contribution in tungsten
is at a minimum when the tungsten layer is farthest from the core; however, as more material is
placed between the core and the tungsten layer, the tungsten layer is in turn positioned further
from the core, increasing volume and mass approximately proportional to the square of its
distance from the reactor. The composite shield configuration was designed to first, maintain
acceptable radiation levels from reaching the payload, and second to minimize mass. As a result,
the lightest and most effective shielding configuration yielded a total mass of 3685 kg (2629 kg
tungsten, 1055 kg lithium hydride).
Reactor Results
In conclusion, we started the design process by simply allocating and tabulating power
requirements for anything and everything that would be on board the CTV, and after several
power systems trade studies, we found that a nuclear reactor would provide the most effective
and efficient means of satisfying those requirements. We illustrate the complete design of the
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reactor in Figure 5.4.4-2 Crew Transfer Vehicle nuclear power system modelwhich depicts every
major component of the system except for the heat rejecting radiator panels.
Figure 5.4.4-2 Crew Transfer Vehicle nuclear power system model
The reactor was optimized in every aspect of its design, from its shielding orientation to its
heat rejection and Power Management and Distribution (PMAD) system to output the lightest
and therefore cheapest, in terms of cost per kilogram, power system available. Below is a table of
the fission reactor’s masses broken down into the most important components just to show and
summarize where the reactor’s masses are all allocated.
Table 1 Reactor Mass Distribution
Power,
kW
Total Mass,
kg
PMAD,
kg
Heat
Rejection,
kg
Power
Conversion,
kg
HSHX,
kg
Shield,
kg
Reactor,
kg
Specific
Power,
kg/kW
2100 25872 5715 3723 5942 2755 3685 4051 12.86
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References
[1] Craft, A.E., and King, J.C., ―Radiation Shielding Options for the Affordable Fission
Surface Power System,‖ Space, Propulsion & Energy Sciences International Forum.
[2] Poston, D.A., Kapernick, R.J., Dixon, D.D., Amiri, B.W., and Marcille, T.F., ―Reference
Reactor Module for the Affordable Fission Surface Power System,‖ Space Technology and
Applications International Forum.
[3] Schmitz, P.C., Schreiber, J.G., and Penswick, L.B., ―Feasibility Study of a Nuclear-Stirling
Power Plant for the Jupiter Icy Moons Orbiter,‖ Space Technology and Applications
International forum.
[4] Houts, M., Hrbud, I., Martin, J., Williams, E., Poston, D., Lipinskit, R., and Ring, P., ―The
Safe Affordable Fission Engine (SAFE) Test Series,‖ NASA/JPL/MSFC/UAH 12th
Annual Advanced Space Propulsion Workshop.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 196
Author: Michael Hill
5.4.5. Propulsion Systems
We design all of the propulsion systems in accordance with Space Propulsion Analysis and
Design. We also employ third party software, Rocket Propulsion Analysis (RPA), written by the
German software engineer Alexander Ponomarenko as an easier-to-use and more flexible version
of NASA’s Chemical Equilibrium with Applications (CEA) software [1]. We input a given
input of chamber pressure, Pc, fuels and oxidizers (and their respective relative weights), desired
thrust, F, expansion area ratio, ε, general nozzle shape (conical or parabolic bell), and any
throttling requirements. The software calculates parameters such as specific impulse, Isp,
vacuum specific impulse, Ispv, characteristic velocity, c*, specific heat ratio, , thrust coefficient,
cf, as well as thermal transport properties. The software also generates the general nozzle
contour that provides the desired thrust using standard contour equations for parabolic and
conical nozzles. Once the software calculates the contours, we calculate masses and volumes of
engines.
5.4.5.1. High Thrust Primary Engines
We use six high-thrust liquid hydrogen (LH2) and liquid oxygen (LOX) engines on the crew
transfer vehicle. Each engine outputs 1.5 million N of thrust at a specific impulse of 458.30
seconds based off of a theoretical vacuum specific impulse of 471.80 seconds and reaction
efficiency of 0.9935 and nozzle efficiency of 0.9777. The engines have an expansion area ratio,
ε, of 120 and a chamber pressure, Pc, of 10 MPa; values based on historical data of space engines
of similar thrust. Using these values, the chemical equation software tells us the engines require
an oxidizer to fuel ratio, O/F, of 5.136 to achieve the optimal specific impulse and a total mass
flow rate of 333.747 kg/s to achieve the required thrust. The engines have a parabolic bell nozzle
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 197
Author: Michael Hill
to decrease the overall required length with a characteristic length, L*, of 1 meter. The following
is the required contour to achieve the desired performance.
Figure 5.4.5.1-3 General nozzle contour geometry of a chemical engine
Table 5.4.5.1-2 Contour geometry parameter values for high-thrust kick engine
(Refer to Fig. 5.4.5.1-1)
Parameter Value Units
Radius of Chamber, Rc 0.268545 m
Chamber Angle, b 30.00 deg
R2 0.28445 m
R1 0.24022 m
Radius of Throat, Rt 0.160145 m
Nozzle Radius, Rn 0.06117 m
Radius of Exit, Re 1.75429 m
Initial Nozzle Angle, Tn 37.43 deg
Final Nozzle Angle, Te 8.29 deg
Length of Cylindrical Portion of Chamber, Lcyl 0.13631 m
Total Length of Combustion Chamber, Lc 0.46465 m
Length of Nozzle, Le 4.76597 m
We choose the material columbium/niobium for the high-thrust engine, which has a high
melting point of 2750 K and a density of 8300 kg/m3. These properties make it a typical material
for this kind of application. We calculate the wall thickness in the combustion chamber to be
34.65 mm thick based off of a rough burst pressure calculation with a factor of safety of ―2‖ and
the ultimate tensile strength, Ftu, of 310 MPa [2]. We calculate the required wall thickness at the
throat to be half of that of the combustion chamber wall thickness, 17.325 mm. We choose a
linear taper of wall thickness from the throat to the exit of the nozzle with the exit of the nozzle
By: Michael Hill
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 198
Author: Michael Hill
at 0 mm to calculate the total mass of the complete combustion chamber and nozzle system to be
643.01 kg.
The kick engine requires a significantly large tank system, so a linearly pressurized system is
not an option. Therefore, we must design turbomachinery to increase the fuel and oxidizer
pressures up to the chamber pressure and account for pressure losses along the way. We assume
a pump efficiency, ηp, of 0.80 and a turbine efficiency, ηt, of 0.70 with a turbine pressure ratio of
8.0. The maximum pressure rise over a single stage of liquid hydrogen is 16 MPa whereas the
maximum pressure rise over a single stage of liquid oxygen is 47 MPa [2]. The maximum
pressure rise per stage is the fundamental factor that determines the number of stages required to
increase pressures levels. Pressure losses in the liquid hydrogen system include coolant system
losses, dynamic losses, feed system losses and injector losses. These pressure losses are tabulated
below.
Table 5.4.5.1-3 Pressure losses for the high-thrust kick engines
System LH2 Pressure Loss (kPa) LOX Pressure Loss (kPa)
Coolant System Pressure Loss 1500 0
Dynamic Pressure Loss 3.549 57.05
Feed System Pressure Loss 35 35
Injector Pressure Loss 2000 2000
We choose tank pressures based on a logarithmic curve fit of historical data based on tank
volume for turbopump-fed systems. We calculate the pressure for the liquid hydrogen tanks and
liquid oxygen tanks to be 233 kPa and 373 kPa, respectively. Since the liquid hydrogen and
liquid oxygen pumps must overcome a total change of pressure of 9.767 MPa and 9.627 Mpa,
respectively, these total changes of pressure are lower than their single-stage maximum pressure
rises and only one stage is required for each pump. The mass of each pump is a function of pump
shaft torque [2]. The overall mass of the turbomachinery is 47.64 kg. We base the masses for the
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 199
Author: Michael Hill
injector, feed system and cooling system off of historical data. Typically, the injector mass is
62.25% of the combustion chamber and nozzle system. The mass of the engine structural
(combustion chamber, nozzle and injector) accounts for 40% of the mass of the entire engine
sans turbomachinery (system mass). The feed system accounts for 24.9% of the system mass
and the cooling system accounts for 35.1% of the system mass [2]. The overall mass of the
engine is the sum of the system mass and the turbomachinery mass. Individual component
masses are tabulated below.
Table 5.4.5.1-4 High-thrust kick engine component masses
System Mass/Engine (kg)
Combustion Chamber 298.21
Nozzle 344.81
Injector 400.28
Feed 649.45
Cooling 915.49
O2 Turbomachinery 42.45
H2 Turbomachinery 5.19
TOTAL 2655.88
Using the calculated total mass, we calculate the high-thrust kick engine to have a thrust to
weight ratio of 57.57. As a sanity check, we compare the predicted performance of our engine to
that of a similar engine, the J-2X. The J-2X is a 1.31 kN liquid hydrogen/liquid oxygen engine
with a thrust to weight ratio of 55.04, dry mass of 2472 kg and an Isp of 448 sec. Our engine
matches very nicely to the J-2X being a slightly higher thrust engine with slightly higher mass.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 200
Author: Michael Hill
5.4.5.2. Electric Low Thrust Motors
We choose to use four magnetoplasmadynamics (MPD) thrusters as electric low thrust motors
to provide thrust for the outbound and return spiral trips as well as to maintain stability. We place
the MPDs at the center of mass on the Crew Transfer Vehicle on an ―Ion Rail‖ which allows the
thrusters to move while firing to maintain true to the center of mass while it is shifting due to
variation in mass distribution in the fuel tanks due to firing. Each thruster is capable of producing
10 N of thrust at 5000 seconds of specific impulse amounting to a total of 40 N of thrust.
The MPDs operate using gaseous hydrogen assumed efficiency of 50%. To operate at these
conditions, we require 490 kW of power per thruster, or 1.96 MW of total power required [3].
We assume the masses of each thruster follow the trend outlined by McGuire – that the thruster
and power processing unit scales by a factor of 1.3552 kg/kWe with the power processing unit
scaling by a factor of 1.25 kg/kWe [4]. Thus, the mass of each engine is 51.5 kg and the mass of
each power processing unit is 664 kg with a combined mass of 715.5 kg. The total mass of the
low-thrust system is 2862 kg.
We choose the MPD system as the result of a trade study with an experimental 30 kW arcjet
reactor since fuel must be produced in-situ; hydrogen and oxygen are our only options as a
working fluid. As a result of the trade study, it is found that the choice of engine relies solely on
the mass of the power generation system. For heavier power systems, the MPD becomes more
attractive than the arcjet, which is our case.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 201
Author: Michael Hill
5.4.5.3. Ceres Regime Engines
The Ceres regime engines provide landing and hovering capabilities on Ceres. Three engines
are present on the crew transfer vehicle; one on each primary tank. They share the same tanks as
the high-thrust kick engines. Like the high-thrust kick engines, the Ceres regime engines have a
parabolic bell nozzle with a characteristic length of 1 meter. As with the kick engine, the
following table outlines the contour geometry of the Ceres regime engine.
Table 5.4.5.3-1 Contour geometry parameter values for Ceres Regime engine
(Refer to Fig. 5.4.5.1-1.)
Parameter Value Units
Radius of Chamber, Rc 0.09844 m
Chamber Angle, b 30.00 deg
R2 0.21735 m
R1 0.05021 m
Radius of Throat, Rt 0.033475 m
Nozzle Radius, Rn 0.01279 m
Radius of Exit, Re 0.33474 m
Initial Nozzle Angle, Tn 36.74 deg
Final Nozzle Angle, Te 8.51 deg
Length of Cylindrical Portion of Chamber, Lcyl 0.03674 m
Total Length of Combustion Chamber, Lc 0.22106 m
Length of Nozzle, Le 0.90082 m
The Ceres regime engine achieves 33.33 kN of thrust at 100% thrust level and throttles down
to 50% thrust. With an optimal oxidizer to fuel mixture ratio to maximize specific impulse is
4.989 with a total mass flow of 70.51 kg/s per engine operating at 33.33 kN. At 100% thrust, we
expect the engine to deliver a specific impulse of 469.08 sec and 468.87 sec at 50% thrust. The
engine achieves a thrust to weight ratio of 61.49, which compares well with other LH2/LOX
engines.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 202
Author: Michael Hill
Co-Author: Evan Helmeid
We size the Ceres regime engines to accommodate two different phases:
Initial optimal descent upon Ceres arrival, nearly empty of propellant
Switching from ISPP 1 to ISPP2 at midterm of Ceres stay
Additionally, the engines must throttle down to provide hover capability upon landing. These
maneuvers require an engine capable of a thrust to weight ratio (T:W) between ~0.7 and ~2.5 at
all times (taking into account the varying mass). As such, the engines yield a total nominal thrust
of 100 kN, and throttle down to a minimum of 10 kN. We outline maneuvers and thrust levels in
Table 5.4.5.3-2.
Table 5.4.5.3-2 Required engine characteristics, boundaries, and actual engine capabilities
Initial Descent ISPP Switch
Required Achievable Required Achievable
Thrust – Nominal [kN] - - 100.0 - - 100.0
Thrust – Min [kN] - - 10.00 - - 10.00
Weight – Max[kN] 45.64 - - 50.57 - -
Weight – Min [kN] 41.95 - - 44.02 - -
T:W Optimal Descent >2.0 2.2~2.4 >2.0 2.0~2.3
T:W Land <0.7 0.23~1.0 <0.7 0.23~1.0
Since we choose to share tank systems with the high-thrust kick engine, we must match the
tank pressure. We also choose to lower the chamber pressure for this engine because of its
smaller size to 5 MPa. Pressure losses and component mass breakdown for the Ceres regime
engine are calculated in the same manner as the high-thrust kick engine and are tabulated below.
Table 5.4.5.3-3 Pressure losses for the Ceres regime engine
System LH2 Pressure Loss
(kPa)
LOX Pressure Loss
(kPa)
Coolant System Pressure Loss 750 0
Dynamic Pressure Loss 3.549 57.05
Feed System Pressure Loss 35 35
Injector Pressure Loss 1000 1000
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 203
Author: Michael Hill
Co-Author: Evan Helmeid
Table 5.4.5.3-4 Ceres regime engine component masses
System Mass/Engine (kg)
Combustion Chamber 9.67
Nozzle 2.50
Injector 7.57
Feed System 12.28
Cooling System 17.32
O2 Turbomachinery 1.28
H2 Turbomachinery 4.66
TOTAL 55.27
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 204
Author: Alex Kreul
Co-Author: Michael Hill
5.4.5.4. Primary Tank System
We design our Crew Transfer Vehicle (CTV) Primary Tank System to hold a portion of the
propellant required for the trip from Earth to Ceres, with the Earth Departure Tank System
holding the rest of the required propellant. At the same time, the Primary Tank System must be
large enough to hold all propellant required for the return trip from Ceres to Earth. These two
tank systems are sized together to maximize the size of the Earth Departure Tank System, and
minimize the size of the Primary Tank System, as detailed in section 5.4.5.5. We size the
Primary Tanks to contain propellant needed by the high-thrust kick engines, the low-thrust MPD
engines, and the Ceres Regime motors.
Figure 5.4.5.4-1 The Primary Tank set. The LH2 tank is the larger tank on top, and the
LOX tank is the smaller tank beneath. Also shown are the kick motor, the Ceres regime
motor, and a landing leg.
By: Alex Roth
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 205
Author: Alex Kreul
Co-Author: Michael Hill
The Primary Tank System consists of three sets of tanks. Each set has a carbon fiber
reinforced plastic (CFRP) liquid hydrogen tank with ellipsoidal end caps, a CFRP spherical
liquid oxygen tank, a kick motor, and a Ceres regime motor. All sets are positioned on the
counterweight portion of the CTV, and each set is identical in size. (Dimensions can be found in
the table below.) Finally, for thermal insulation and micrometeoroid protection purposes, we
cover each tank with 50 layers of multi-layer insulation (MLI) with a thickness of 0.0125 m.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 206
Author: Alex Kreul
Co-Author: Michael Hill
Table 5.4.5.4-1 Primary Tank System specifications. This table includes the LH2 tank
dimensions, as well as mass and volume for individual tanks and all three tanks combined.
Cylinder Elliptical caps
Length, m 3.337
Inner major radius, m 4.38
Inner radius, m 4.38
Inner minor radius, m 2.19
Vessel thickness, 0.0023 Vessel thickness, m 0.0023
Total length, m 7.746 Volume of 1, m
3 380
Volume of 3, m3 1139
Mass of 1, kg 884 Mass of 3, kg 2653
Table 5.4.5.4-2 Primary Tank System specifications. This table includes the LOX tank
dimensions, as well as mass and volume for individual tanks and all three tanks combined.
Sphere
Inner radius, m 2.802
Vessel thickness, m 0.0032
Total length, m 5.636
Volume of 1, m3 93
Volume of 3, m3 280
Mass of 1, kg 487
Mass of 3, kg 1460
Both the high-thrust kick engines and the Ceres regime engines are autogenously pressurized
systems, meaning the tanks are pressurized with hot gasses resulting from cooling the engine.
We choose to use hydrogen as the cooling fluid because corrosion becomes problematic [2]. A
pump draws liquid hydrogen into the cooling jacket of the thruster to become gaseous. The
gaseous hydrogen goes through a turbine, which mechanically powers the liquid hydrogen and
liquid oxygen pump. Oxygen is drawn off into a bypass system to interface with the gaseous
hydrogen in a heat exchanger to vaporize. This oxygen vapor now becomes the pressurant for
the liquid oxygen tank. Once the kick motors are finished firing, the liquid hydrogen tanks have
now filled with excess gaseous hydrogen. This gaseous hydrogen is then used to fuel the MPD
thrusters. We do this to decrease the number of required tanks and so that we do not launch
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 207
Author: Alex Kreul
Co-Author: Michael Hill
gaseous hydrogen from Earth, which would be a waste of space. Fig. 5.4.5.4-44-2 is a graphical
representation of how the primary tank system works.
Fig. 5.4.5.4-4. Autogenously cooled LH2/LOX thruster and tank system
The Earth departure tank system consists of three tanks similar to the primary tank system,
but tanks are jettisoned after the main burn at Earth to reduce weight. They are smaller in size
because they do not need to carry fuel for a burn at Ceres or to provide fuel for the Ceres regime
motors.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 208
Author: Alex Kreul
Co-Author: Michael Hill
5.4.5.5 Earth Departure Tank System
We design the Earth Departure Tank System to hold the remainder of the required propellant
for the trip from Earth to Ceres. Upon arrival at Ceres, we discard the tank system to reduce the
Crew Transfer Vehicle (CTV) mass for the return trip. We size the tank system to maximize the
propellant mass without exceeding the mass restrictions of the Ares V, our heavy lift vehicle.
Figure 5.4.5.5-1 The Earth Departure Tank set. The LH2tank is the larger tank on top,
and the LOX tank is the smaller tank beneath. Also shown is the kick motor.
The Earth Departure Tank System consists of three sets of tanks. Each set has a carbon fiber
reinforced plastic (CFRP) cylindrical liquid hydrogen tank with ellipsoidal end caps, a CFRP
spherical liquid oxygen tank, and a kick motor. All sets are positioned on the counterweight
portion of the CTV, and each set is identical in size, which is larger than the primary tank system
sets. (Dimensions can be found in the table below.) Finally, for thermal insulation and
micrometeoroid protection purposes, we cover each tank with 50 layers of multi-layer insulation
(MLI) with a thickness of 0.0125 m.
By: Alex Roth
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 209
Author: Alex Kreul
Co-Author: Michael Hill
Table 5.4.5.5-1 Earth Departure Tank System specifications. This table includes the LH2
tank dimensions, as well as mass and volume for individual tanks and all three tanks
combined.
Cylinder Elliptical caps
Length, m 6.634
Inner major radius, m 4.38
Inner radius, m 4.38
Inner minor radius, m 2.19
Vessel thickness, m 0.0023 Vessel thickness, m 0.0023
Total length, m 10.044 Volume of 1, m
3 519
Volume of 3, m3 1557
Mass of 1, kg 1106 Mass of 3, kg 3317
Table 5.4.5.5-2 Earth Departure Tank System specifications. This table includes the LOX
tank dimensions, as well as mass and volume for individual tanks and all three tanks
combined.
Sphere
Inner radius, m 3.251
Vessel thickness, m 0.0039
Total length, m 6.5354
Volume of 1, m3 145
Volume of 3, m3 436
Mass of 1, kg 806
Mass of 3, kg 2419
5.4.5.6. Final Orbit Raise and Stabilization Engine (FORSE)
The FORSE engine is exactly the same engine used for the Ceres Regime motors. The only
difference is location on the Crew Transfer Vehicle and the fact it has its own tank system
because it serves a different purpose than the Ceres regime motors. Thus, geometry,
pressure losses and component masses are shared with the Ceres regime motors in
Table 5.4.5.3-, are tabulated below.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 210
Author: Michael Hill
Co-Author: Evan Helmeid
Table 5.4.5.3-, and Table 5.4.5.3-.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 211
Author: Michael Hill
Co-Author: Evan Helmeid
5.4.5.7 FORSE Tank System
We design the FORSE tank system is to hold the propellant required for Crew Transfer
Vehicle (CTV) Earth re-entry maneuvers, once the primary tank system has been discarded. The
system consists of one set of tanks, including a carbon fiber reinforced plastic (CFRP) spherical
liquid hydrogen tank and a CFRP spherical liquid oxygen tank. This set of tanks is positioned on
the crew quarters and storage attic. (Dimensions can be found in the table below.) Finally, for
thermal insulation and micrometeoroid protection purposes, each tank is covered with 50 layers
of multi-layer insulation (MLI) with a thickness of 0.0125 m.
Table 5.4.5.7-1 FORSE Tank System specifications. This table includes the LH2 tank
dimensions, mass, and volume.
Sphere
Inner radius, m 1.155
Vessel thickness, m 0.002
Total length, m 2.314
Volume, m3 6.62
Mass, kg 51.89
Table 5.4.5.7-2 FORSE Tank System specifications. This table includes the LOX tank
dimensions, mass, and volume.
Sphere
Inner radius, m 0.783
Vessel thickness, m 0.002
Total length, m 1.560
Volume, m3 2.09
Mass, kg 23.86
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 212
Author: Alex Kreul
Co-Author: Michael Hill
5.4.5.8. Attitude Thrusters
We choose to use three hypergolic attitude thrusters. We desire to provide 30 N of thrust per
engine. We achieve this by scaling up a Pratt & Whitney Rocketdyne RS-43 engine. We are
able to provide thrust at 284 seconds of Isp. Each engine has a mass of 0.93 kg resulting in a total
mass of 2.79 kg assuming mass linearly scales with thrust requirement.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 213
Author: Alex Kreul
Co-Author: Michael Hill
5.4.5.9. Attitude Thruster Tank System
We design the attitude thruster tank system to hold the propellant required for attitude control
of the Crew Transfer Vehicle (CTV). The system consists of six sets of tanks. Each set has a
carbon fiber reinforced plastic (CFRP) spherical liquid hydrogen tank and a CFRP spherical
liquid oxygen tank. Three sets are positioned on both the crew portion of the CTV and the
counterweight portion of the CTV. Each set is identical in size. (Dimensions can be found in the
table below.) Finally, for thermal insulation and micrometeoroid protection purposes, we cover
each tank with 50 layers of multi-layer insulation (MLI) with a thickness of 0.0125 m.
Table 5.4.5.4-1 Attitude Thruster Tank System specifications. This table includes the LH2
tank dimensions, mass, and volume.
Sphere
Inner radius, m 0.223
Vessel thickness, m 0.002
Total length, m 0.450
Volume of 1, m3 0.22
Volume of 6, m3 1.34
Mass of 1, kg 5.27
Mass of 6, kg 31.59
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 214
Author: Alex Kreul
Co-Author: Michael Hill
Table 5.4.5.4-2 Attitude Thruster Tank System specifications. This table includes the LOX
tank dimensions, mass, and volume.
Sphere
Inner radius, m 0.203
Vessel thickness, m 0.002
Total length, m 0.410
Volume of 1, m3 0.17
Volume of 6, m3 1.03
Mass of 1, kg 4.40
Mass of 6, kg 26.39
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 215
Author: Alex Kreul
Co-Author: Michael Hill
References
[1] Ponomarenko, A., ―RPA: Tool for Liquid Propellant Rocket Engine Analysis C++
Implementation,‖ May 2010.
[2] Humble, R.W., Henry, G.N., Larson, W.J., ―Liquid Rocket Propulsion Systems,‖ Space
Propulsion Analysis and Design, 1st ed. (rev.), Space Technology Series, McGraw-Hill,
New York, 1995, pp. 179-294.
[3] Humble, R.W., Henry, G.N., Larson, W.J., ―Electric Rocket Propulsion Systems,‖ Space
Propulsion Analysis and Design, 1st ed. (rev.), Space Technology Series, McGraw-Hill,
New York, 1995, pp. 509-598.
[4] McGuire, Melissa L., Borowski, Stanley K., Mason, Lee M., ―High Power MPD Nuclear
Electric Propulsion (NEP) for Artificial Gravity HOPE Missions to Callisto.‖ NASA/TM-
2003-212349, Dec. 2003.
Detailed Vehicle Designs Crew Transfer Vehicle Page 216
Author: Chris Luken
Co-Author: Frank Fortunato
5.4.6 Attitude Determination and Control Systems
During all Crew Transfer Vehicle (CTV) operational stages, we maintain proper orientation
and configuration with a slew of control systems and devices. Due to our implementation of
tethers to simulate an artificial Martian gravity, we require new methods of stabilization and
orientation control. Control devices ranging from innovative attitude control thrusters, to an
unconventional tether system make the CTV a whole new breed of spacecraft. We see from Fig.
5.4.6-1 that the spacecraft extends with its unique tether design. Despite higher complexity, this
feature reduces structural mass and allows maximum flexibility for the simulation of artificial
gravity at any level. The functionality of our flexible vehicle design changes vehicle inertia
characteristics and the magnitude of artificial gravity. We must consider various aspects and
control systems of the CTV for analysis, including measurement device accuracy levels, natural
frequencies, and the rigid body assumption.
Figure 5.4.6-1 Extended CTV configuration displaying the vehicle sections and tethers
The following sub-sections investigate the necessary stabilization systems for the CTV,
required vehicle maneuvers during orbit transfers for burn alignments, simulated gravity, and
Alex Roth
Detailed Vehicle Designs Crew Transfer Vehicle Page 217
Author: Chris Luken
Co-Author: Frank Fortunato
methods implemented to turn the all-spin vehicle around the sun. As a part of the vehicle turning,
an innovative design combining the tether system and the low thrust motors into a multi-use
device dubbed the ―IonRail.‖
Table 5.4.6-1 below provides a full breakdown of attitude propellant requirements for both
outbound and return trips. The propellant stabilizes the vehicle due to perturbations from
environmental torques and forces, vehicle re-orientation maneuvers, and spin-up/de-spin events.
Table 5.4.6-1 Attitude propellant masses for Crew Transfer Vehicle
Outbound, kg Ceres Operations, kg Return, kg
Environmental Torques 1,973 -- 1,958
Spin-up/De-spin Events 758 -- 613
Re-orientation Maneuvers 53 -- 26.6
Ceres Operations 78.5 551 50.1
Total Propellant 2,863 551 2,649
Attitude Thrusters 5.6 5.6 5.6
Alternate Control Devices 64 64 64
Total Attitude Hardware 69.6 69.6 69.6
We discuss each of the elements in the table above in the following sections, focusing on
stabilization (5.4.6.1), vehicle re-orientation (5.4.6.2), artificial gravity (5.4.6.3), and all-spin
turning maneuvers (5.4.6.4). The Ceres operations masses account for additional torque due to
non-zero products of inertia when the capsule is repositioned the crew habitat side of the vehicle.
We position the capsule to allow for reasonable loading of regolith from Ceres.
We must also account for the variety of control device masses. They consist of six thrusters
mounted on Canfield joints. The alternate control devices we implement on the CTV include
CMG’s, tether actuators, and dampers on both ends of the vehicle. We implement dampers to
help eliminate angular momentum about the vehicle axis of symmetry.
Detailed Vehicle Designs Crew Transfer Vehicle Page 218
Author: Chris Luken
5.4.6.1 Stabilization
During compact mode, wherein the vehicle sections are fully retracted, we provide general
stabilization with six attitude thrusters for large corrections and Control Moment Gyros (CMGs)
for minute adjustments. As we extend the vehicle for artificial gravity, attitude thrusters provide
the moments required to induce spin. We also use these devices to counteract environmental
torques acting on the body, including the effects of gravity from massive gravitational bodies,
solar radiation, solar wind and micrometeorites.
We install attitude measurement devices such as star sensors, sun sensors, accelerometers, and
gyroscopes on the CTV to provide accurate position and orientation data. Control systems will
be developed that use combinations of available devices including attitude thrusters, CMGs,
tether actuators, and the ―IonRail‖ system. Also, we use dampers as described in the attitude
control section (5.4.6) to eliminate angular momentum in unwanted directions. These systems
handle perturbations due to environmental torques, variation in inertia characteristics due to crew
movement or fuel sloshing, spin-up and de-spin maneuvers, vehicle orientation changes, and
turning the vehicle around the Sun.
5.4.6.1.1 Thruster Positioning
We position six attitude thrusters on the vehicle as shown in Fig. 5.4.6.1.1-2. The 50 N
thrusters are mounted on specialized actuator joints known as Canfield joints. These joints are
currently under development by Professor Canfield at Tennessee Technological University. We
position three thrusters at the aft end of the tank and reactor assembly with the remaining three
on the far end of the crew habitat section. These are uniformly arrayed on a circle as shown in
Fig. 5.4.6.1.1-3.
Detailed Vehicle Designs Crew Transfer Vehicle Page 219
Author: Chris Luken
Figure 5.4.6.1.1-2 Thruster positioning on the extended CTV mode (exaggerated size)
Attitude thrusters not to scale.
The figure above indicates the positioning of 50 N attitude thrusters on the CTV, connected to
the main chassis. Based on the computed tension in the tether during the all-spin extended mode,
we can apply the rigid body assumption to further analysis.
Figure 5.4.6.1.1-3 General thruster arrangement viewed from the aft or front end
Viewing the CTV from either end, we array the thrusters 120 degrees apart. The Canfield
joints on which the thrusters are mounted allow us to point two thrusters in any given direction.
This arrangement gives maximum efficiency when applying a torque about the vehicle center of.
Christopher Luken
Detailed Vehicle Designs Crew Transfer Vehicle Page 220
Author: Chris Luken
5.4.6.1.2. Environmental Torques
A survey of higher order environmental torques must be assessed and an appropriate level of
attitude propellant allotted to counter said torques. By several orders of magnitude, the largest
environmental torque is gravitational, supplied by the Sun and Jupiter during the transfer from
Earth to Ceres and back. Following these sources of environmental torque are those of solar
radiation, solar wind, and micrometeorites. Due to the very large inertia characteristics of the
vehicle and the large surface areas provided by the vehicle tanks and crew habitation module, the
environmental torques acting on the vehicle are substantial.
Before jumping into an analysis of these four acting torques, we want to determine the vehicle
characteristics and define the constants used in this assessment. These values are shown in Table
5.4.6.1.2-1 below.
Table 5.4.6.1.2-1 Crew Transfer Vehicle basic characteristics
Characteristic Value Units
Vehicle Mass (Outbound) 333,078 kg
Vehicle Mass (Return) 189,107 kg
Transverse Inertia (Outbound) 7.248x108 kg-m
2
Axial Inertia (Outbound) 3.876x106 kg-m
2
Transverse Inertia (Return) 1.029x109 kg-m
2
Axial Inertia (Return) 1.455x106 kg-m
2
Maximum Surface Area (Extended) 1500 m2
Minimum Surface Area (Extended) 500 m2
µ Sun 1.3272x1020
N-m2/kg
µ Jupiter 1.267x1017
N-m2/kg
With these values, we are now ready to assess the acting environmental torques on the CTV
and determine an appropriate amount of attitude fuel to counteract their resulting moments. It is
important to note that all environmental torques excluding micrometeorites are subject to the
Detailed Vehicle Designs Crew Transfer Vehicle Page 221
Author: Chris Luken
inverse-square law. This means that the further the vehicle is from the source, the resulting
torques will decrease.
Gravitational Torques
As noted, the largest environmental torque is gravitational. For the analysis carried out in the
appendix, the distance from the sun is computed via numerical integration of the heliocentric
transfer orbits. To consider the worst case scenario for gravitational torque contribution, Jupiter
is presumed to be directly in line with the Sun and CTV.
Since the Vehicle is spinning about an axis which is parallel to the orbit fixed radial plane, the
magnitude of the torque due to the sun oscillates as the vehicle spins. Figure 5.4.6.1.2-4 below
provides graphical representation of the spinning orientation of the vehicle relative to the Sun
and Jupiter.
Figure 5.4.6.1.2-4 CTV orientation with respect to the Sun and Jupiter and reference
frames
For each revolution, there are two periods of torque oscillation since the vehicle is spinning
about a transverse axis. While the vehicle is spinning, the sinusoidal solar moments peak at Earth
Detailed Vehicle Designs Crew Transfer Vehicle Page 222
Author: Chris Luken
and significantly decrease during the trip. Over the course of one revolution, the net moment
effectively cancels, however the magnitude near earth provides a source of concern as it could
potentially de-spin the extended vehicle. If this happens, the rigid body assumption no longer
holds and the vehicle will collapse. To counter this issue, we can allot an amount of fuel to
provide necessary counter moments which will ensure that the vehicle remains spinning. Table
5.4.6.1.2-2 provides maximum moments at three discrete times during the transfer out to Ceres
and the return transfer to Earth.
Table 5.4.6.1.2-2 Peak gravitational torques on CTV
Distance From Sun Peak Torque, kNm
Earth (1.0 AU) 42.7
Outbound Midpoint (1.385 AU) 16.1
Ceres (3.2 AU) 2.48
Return Midpoint(2.658 AU) 4.236
Return Earth (1.035 AU) 48.99
Outbound Propellant Requirement 1938 kg
Return Propellant Requirement 1924 kg
Comparing the peak torques to the tension in the tether system connecting the two vehicles
indicates that the rigid body assumption can be maintained. The maximum gravitational torque
caused by the Sun on the CTV is smaller by 2 orders of magnitude. Based on this, earlier
concerns about the body losing rigidity are moot.
Solar Radiation
Solar Radiation is the next most significant environmental torque, however compared to the
gravitational torques; it is several orders of magnitude smaller than gravitational. For the case of
the CTV, we need to first find the area centroids of both sides of the vehicle. The distance of
Detailed Vehicle Designs Crew Transfer Vehicle Page 223
Author: Chris Luken
these centroids from the vehicle center of mass provide us with the acting moment arms for solar
radiation.
As with the gravity moment analysis, the all spin vehicle as shown in Fig. A-5.4.6.1.2-2 must
deal with significant variations in torque. While the source nature of this torque is completely
different to the gravitational torque previously investigated, the variation during each revolution
of the vehicle is the same. Before diving into torque calculations however, it is a good idea to
first check the magnitude of the acting forces on the CTV. An expression for the magnitude of
solar radiation forces is shown below in equation 5.4.6.1.2-1. [1]
(5.4.6.1.2-1)
A description of this expression is provided in the appendix correlating to this section of the
report. The constants used, and the correlating force values acting on the CTV as a function of
vehicle area at time t yield a force magnitude of 9.02x10-6
A(t) N.
It is readily apparent that the force is minimal and we can assume that onboard CMG’s can
account for these forces. The areas of either side of the CTV do not exceed one million square
meters, so the likelihood of attitude correction is unnecessary, and as such, we do not need to
compute the resulting moments. It is still important to account for the velocity change that these
forces will impact with the vehicle. Using Newton’s 2nd
law, the spacecraft mass, and the attitude
thruster specific impulse along with the transfer duration, the required correctional propellant is
easily determined with the rocket equation. Table 5.4.6.1.2-3 provides a breakdown of the values
implemented.
Detailed Vehicle Designs Crew Transfer Vehicle Page 224
Author: Chris Luken
Table 5.4.6.1.2-3 Correctional attitude propellant required for Crew Transfer Vehicle
Variable of Interest Value Units
CTV Mass (outbound) 333,078 kg
CTV Mass (return) 189,107 kg
Thruster Isp (MMH/NTO) 328 s
Outbound Propellant Required 34.2 kg
Return Propellant Required 33.9 kg
Solar Wind
Similar in nature to solar radiation, solar wind provides a moment and force on the CTV.
With an expression detailed in the appendix covering the CTV environmental torques, the
maximum force turns out to be 2.3x10-9
A(t) N. Thus the resulting moment contributions to the
vehicle can be considered to be negligible.
Micro-meteoroids
Micro-meteoroids, initially investigated by Donald Kessler provide a significant hazard to any
spacecraft. Apart from potentially ripping holes through space vessels, they can also provide a
moment about the CTV center of mass. A quick investigation provided in the appendix yields a
maximum force of roughly 8x10-10
A(t) N exists. This was computed based on the average
meteoroid density in space at the asteroid belt wherein Ceres exists which is assumed to be the
highest density during the journey. Overall, the contribution can be considered negligible.
From a mission safety point of view, it is suggested that Whipple Shields be implemented
which disintegrate micro-meteoroids as they pass through a material held away from the vehicle
hull.
Detailed Vehicle Designs Crew Transfer Vehicle Page 225
Author: Chris Luken
References
[1] Longuski, J.M.,Todd,R.E.,Konig,W.W.,‖Survey of Nongravitational Forces and Space
Environmental Torques: Applied to the Galileo‖, Journal of Guidance, Control, and
Dynamics, Col. 15, No. 3, pp 545-553
Detailed Vehicle Designs Crew Transfer Vehicle Page 226
Author: Frank Fortunato
5.4.6.1.3. Tether-Reel System
A powered reel system controls the length and rate of change in cable length. Because certain
phases of the mission require the vehicle to change configuration, the crew habitat extends in and
out of the center chassis using the tether-reel system. The entire crew transfer vehicle spins and
thrusts about a point on the tether. This creates the need for three cables versus a single tether,
which provides bending and torsion stability. With the combination of environmental forces,
vehicle maneuvering, and internal dynamics, the three tethers provide the necessary stability
while under tension. Three individual cables support the crew habitat and each one has their own
pulley system, shown in Fig. 5.4.6.1.3-1.
Figure 5.4.6.1.3-1 Schematic of Crew Section Extended Outward by Tether-Reel System
In the previous figure, the tethers attach on the outside of the storage attic 120 degrees from
each other. This provides the vehicle tether system with axis symmetry and static stability. We
size the tethers to handle ten times the operating tension and 110% maximum operating length
Detailed Vehicle Designs Crew Transfer Vehicle Page 227
Author: Frank Fortunato
discussed in section 5.4.6.3 by Frank Fortunato. This factor of safety provides the vehicle the
ability to withstand conditions beyond the ideal mission specifics.
5.4.6.1.4. Tether Actuators – Crew Habitat Orientation
We put sets of actuators in place to stabilize against vibrations and other undesired behaviors.
Each actuator contains the capability of varying the length of individual cables a small amount.
Another important aspect of this feature is the control of the relative orientation of each
component on each end of the tethers. This provides the CTV with the ability to change the
direction of artificial gravity for the crew quarters. Inside the living space, crewmembers and
other components change location and ultimately affecting the location center of mass on the end
of the tether. Components of the CTV will shift and reorient themselves invalidating the rigid
body assumption. The tether actuators compensate for the relatively small perturbations and
displacements of this kind. Each cable adjusts individually damping vibrations in the tether and
keeping gravity in the desired direction, normal to the floor.
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Author: Frank Fortunato
5.4.6.2. Vehicle Re-Orientation
We can see a timeline of vehicle events, maneuvers, and associated propellant cost in Table
5.4.6.2-1 below. These maneuvers take place before major burns, prior to transfer orbit
initialization, and after transfer orbits completion. We maintain the CTV orientation during
transfer orbits according to the steering law associated with the constant low thrust transfer orbit
described in the CTV trajectory sections (5.4.3 and 5.4.14). This results in three major re-
orientation maneuvers during each transfer, resulting in six major maneuvers throughout the
mission. During each transfer, one re-orientation maneuver is required to align the CTV for
required transfer kick burns. The first of these is a re-orientation from stable Earth orbit
orientation for the outbound trajectory and the other from Ceres for the return trajectory. These
maneuvers also must account for the inclination change between Ceres and Earth. After these
kick burns, we must again re-orient the vehicle to match the initial transfer steering law at the
start of each transfer. When each transfer terminates, we complete a similar maneuver, realigning
the vehicle for orbit insertion. The IonRail, discussed in detail in section 5.4.6.4, maintains the
steering law during each transfer.
We account for expected perturbations with propellant designated as ―General Stabilization‖.
These perturbations include anticipated moments when we position the crew capsule on the side
of the CTV during Ceres operations and anticipated uncertainties directly after ballute separation
during aerocapture.
Detailed Vehicle Designs Crew Transfer Vehicle Page 229
Author: Frank Fortunato
Table 5.4.6.2-1 Crew Transfer Vehicle maneuvers and required propellant
Vehicle Events Maneuver Angle, deg Required Propellant, kg
Crew Boarding (LEO) -- --
Earth Departure Re-Orientation 94.4 27.18
Earth Departure Kick -- --
Heliocentric Transfer Re-Orientation/Spin-up 96.5 12.76
Outbound Heliocentric Transfer -- --
Ceres Arrival Re-Orientation/De-spin 98.4 12.09
Ceres Arrival Kick -- --
Ceres Operations General Stabilization 1290.7
Ceres Return Kick Re-Orientation 94.4 7.43
Ceres Departure Kick -- --
Heliocentric Transfer Re-Orientation/Spin-up 82.9 13.79
Return Heliocentric Transfer -- --
Earth Aerocapture Re-Orientation/De-spin 79.0 5.38
Aerocapture Stabilization/End of Life -- --
Total Propellant Required 1369.33
Detailed Vehicle Designs Crew Transfer Vehicle Page 230
Author: Frank Fortunato
5.4.6.3. Artificial Gravity
For the duration of the transfer from Earth to Ceres and the return trip, the crew spends most
of the mission on board the Crew Transfer Vehicle. We are given the requirement for artificial
gravity for the crew during these stages of the mission. The six crew members experience 0.38
g’s of simulated gravity which is equal to Martian gravity. The CTV changes configuration and
performs a specific maneuver in creating centripetal acceleration. This acceleration results in
artificial gravity for the crew. Ben Stirgwolt of the human factors and science group provides this
spinning maneuver with an operating rate of 2 revolutions per minute discussed in section
5.4.7.11.
The Crew Transfer Vehicle contains the devices and systems capable of performing various
maneuvers which include creating pseudo-gravity for the crew members. The tether system and
the attitude thrusters supply each component of simulated gravity. Accelerating up to the angular
velocity of 2 rpm, attitude thrusters apply the necessary torque on the vehicle. The tethers extend
the crew living quarters away from the rest of the vehicle. This increases the distance between
the center of mass and the crew. The reel system controls the length of the cable during this spin
up process to reduce propellant cost.
The variables of angular acceleration consist of radius and angular velocity. Martian gravity
and 2 rpm equate to a specific crew distance of 84.983 meters. This fixed distance, defined as
one Kwan, normalizes any distance from the center of mass. The Kwan compares length relative
to the necessary operating distance of the center of mass to the crew. The Kwan defines lengths
of the Crew Transfer Vehicle in various configurations. Equation 5.4.6.3-1 on the following page
calculates the radial distance from the vehicle center of mass:
Detailed Vehicle Designs Crew Transfer Vehicle Page 231
Author: Frank Fortunato
(5.4.6.3-1)
Tether extension balances the vehicle and properly locates the spin axis. Propellant tanks and
the chassis structure act as a counter weight for the vehicle, while the remaining mass acts as a
counter weight to the crew living quarters. The counter weight section of the CTV attaches to the
other end of the tether. The counter weight placed on the opposite side of the crew adds to the
total distance. The combination of these distances result in a total tether length greater than 1
Kwan. Tether length adjusts for the change in vehicle mass during the transfer journey between
Earth and Ceres. The center of mass kept at 1 Kwan from the crew maintains the 100 % Martian
gravity specification inside the crew habitat. Figure 5.4.6.3.-1 shows the technique to balance the
vehicle.
Figure 5.4.6.3-1 Rigid body and point mass model of the CTV with the Kwan
Equation 5.4.6.3-2 shows the equality for the mass and distance ratio of the two sections of
the vehicle:
(5.4.6.3-2)
The inputs for computing the main spin-up parameters such as final tether length and total
spin-up time consist of spin rate and gravity. We cycle each unknown through iterations until
Detailed Vehicle Designs Crew Transfer Vehicle Page 232
Author: Frank Fortunato
reaching a solution. The following table shows the values containing results of the artificial
gravity analysis. The spin-up and de-spin maneuvers throughout the mission provide parameters
for calculating the variables. The tethers are sized to handle the maximum operating loads and
distances. Table 5.4.6.3-1 breaks down the size of the tether system for maximum operating
conditions. The largest tether lengths occur at the end of the return trip when the vehicle contains
no propellant. Table 5.4.6.3-2 lists the propellant masses required and maneuver times.
Table 5.4.6.3-1 Tether sizes for the CTV
Parameter Value Unit
Tether Length 178.84 m
2.1 Kwan
Tether Diameter 4.7 cm
Tether Mass 918.2 kg
(Note: 1 Kwan is 84.983 m, the distance required for Martian gravity at 2 rpm)
We integrate the torque over time results in the propellant cost for this maneuver. Calculations
in this analysis use a summation of discrete time steps due to the number of unknowns.
(5.4.6.3-3)
Table 5.4.6.3-2 Crew Transfer Vehicle spin parameters
Parameter Outbound
Spin-Up
Outbound
De-Spin
Return
Spin-Up
Return
De-Spin
Unit
Maneuver Time 2190 2010 1295 1084 s
0.608 0.558 0.359 0.301 hr
Propellant Mass 239.9 268.0 324.1 288.7 kg
Detailed Vehicle Designs Crew Transfer Vehicle Page 233
Author: Frank Fortunato
5.4.6.4. The IonRail
During the transfer orbit between Earth and Ceres we steer the Crew Transfer Vehicle using
the electric motors. Positioning the Magnetoplasmadynamic, MPD, thrusters in the desired
location turns the CTV around the Sun. The mission design group determines our orbit transfer
and flight path angle providing us with the required turning law for the vehicle. We vector the
thrust form the interplanetary motors to follow the determined orbit path. The IonRail system
controls the sliding of the Magnetoplasmadynamic (MPD) thrusters along the tethers for vehicle
turning and navigation.
The motivation for implementing this concept eliminates the large propellant cost from other
systems. We use an already existing system to perform an additional task in addition to the
primary function. The MPD thrusters slide along the tethers to track the vehicle center of mass
during the transfer orbit. The provided thrust from the MPD motors do not apply continuous
torque and keep the vehicle from becoming misaligned. However, in order to turn the vehicle we
desire a certain torque. Shifting the electric thrusters off the center of mass turns the vehicle
when at a specific orientation in its spin.
Space vehicles maneuver and turn easily when there is little to no angular momentum in the
system. In our case, the CTV spins to create artificial gravity as it turns relative to the Sun. The
magnitude of angular momentum must remain constant and only change in direction so that
artificial gravity is constant. Using the IonRail, we move the electric motors out to a distance
creating a moment arm and ultimately a torque on the vehicle.
The IonRail completes several key mission requirements and reduces the cost of other
systems on board the CTV. Most importantly, we apply the interplanetary thrust for the
trajectory at the vehicle center of mass for stability purposes. Second, the MPD thrusters move
Detailed Vehicle Designs Crew Transfer Vehicle Page 234
Author: Frank Fortunato
along the tethers because the center of mass location is not constant. Lastly, the vehicle turns
using an induced torque from positioning the electric motors a distance from the spin axis. This
implemented mechanism creates the same torque on the system by eliminating the use of
chemical thrusters or very large control moment gyroscopes. Figure 5.4.6.4-1 shows the MPD
thrusters sliding on the tethers. By design, the MPD thrusters continuously fire eliminating extra
propellant. Alternate methods require extra propellant expulsion in addition to that of the existing
amount by the low thrust electric motors.
Figure 5.4.6.4-1 The IonRail system sliding along the tethers.
Appendix A.5.4.6.4 outlines the derivation leading to the IonRail equations. The analysis of
the IonRail and its capabilities took torque error and thrust efficiency along with feasibility of
such a device. The magnitude of angular momentum remains constant for artificial gravity
purposes, which provides the IonRail with the primary constraint. The turning law requires that
the direction of angular momentum change for the interplanetary transfer without affecting
By: Alex Roth
Detailed Vehicle Designs Crew Transfer Vehicle Page 235
Author: Frank Fortunato
magnitude. Figure 5.4.6.4-2 depicts the relation of momentum and change in momentum in each
pulse.
Figure 5.4.6.4-2 Schematic showing momentum change by angle psi
Spin conditions and the desired turn angle per half revolution determine the maximum slide
distance for the IonRail. Equation 5.4.6.4-1 shows the driving formula for turning controller.
(5.4.6.4-1)
This provides the IonRail control system with the necessary distance to slide the MPD
thrusters out to and back during the burn arc. Angular velocity and arc angle define the burn arc
time shown below in equation 5.4.6.4-2.
(5.4.6.4-2)
The design of the IonRail provides the turning control system with maximum flexibility. Each
variable parameter combines to perform the turning maneuver with many possible solutions. For
our initial analysis, we determine the burn arc for the IonRail. Figure 5.4.6.4-3 shows the
position trace as a function of burn time for the maximum turn rate.
ψ
Detailed Vehicle Designs Crew Transfer Vehicle Page 236
Author: Frank Fortunato
Figure 5.4.6.4-3 CTV thruster position over one revolution
-30 -20 -10 0 10 20 30
-20
-15
-10
-5
0
5
10
15
20
Max Turn Rate With Variable Burn Arc Angle
IonRail X Position (m)
IonR
ail
Y P
ositio
n (
m)
MPD Thrusters
Max Radius
Burn Arc Boundary
Detailed Vehicle Designs Crew Transfer Vehicle Page 237
Author: Frank Fortunato
Figure 5.4.6.4-4 Close-up of CTV thruster position over one revolution
In Fig. 5.4.6.4-4, we see that the maximum moment arm extending in the vertical direction is
nearly independent of burn arc. 2.14 meters and 1.995 meters comprise the limits of the vertical
displacement. The IonRail sliders create a slight error because of this offset torque.
From the results outlined in appendix A.5.4.6.4, we choose the maximum burn arc for the
IonRail sliders to be 90-degrees. Very small wobble angles and minimal thrust error allow us to
make this decision. Slower moving parts allow the system to operate safely with low chance of
failure. For all of the calculations we use the 90-degree burn arc in operating the IonRail to
minimize slide speed and distance. Table 5.4.6.4-1 shows results using maximum and minimum
inertias to turn at the fastest rate in the mission.
0 2 4 6 8 10 12 14
-4
-2
0
2
4
6
Max Turn Rate With Variable Burn Arc Angle
IonRail X Position (m)
IonR
ail
Y P
ositio
n (
m)
MPD Thrusters
Max Radius
Burn Arc Boundary
Detailed Vehicle Designs Crew Transfer Vehicle Page 238
Author: Frank Fortunato
Table 5.4.6.4-1 IonRail Turning 0.29 Micro-Radians per Second
Moment of Inertia, kg-m2 Force, N Slide Radius, m Slide Speed, m/s
7.142e8 40 2.12 0.44
8.743e8 20 5.19 1.09
The trajectory for the outbound trip uses all of the thrust capability of the MPDs, where as the
half of the maximum 40 N propel the CTV on the return journey. Figure 5.4.6.4-5 shows a plot
of the slide radius versus slider distance.
Figure 5.4.6.4-5 IonRail slider distance for varying turning rates
Vehicle mass and tether length change the vehicle moments of inertia between the two
portions of the mission. The IonRail has the capability to perform the maximum mission turn rate
of 0.29 micro-radians per second. With the vehicle performing the MPD sliding twice per
revolution and the 90 half angle arc there virtually no time where the IonRail is at zero distance.
Figure 5.4.6.4-5 shows traces of the MPD thrust location for various turn rates.
0 0.05 0.1 0.15 0.2 0.25 0.3 0.350
1
2
3
4
5
6
Slider Distance vs. Turn Rate
Turn Rate (rad/s)
Slid
e R
adiu
s (
m)
Detailed Vehicle Designs Crew Transfer Vehicle Page 239
Author: Frank Fortunato
Figure 5.4.6.4-6 IonRail slider location during one revolution
In conclusion, the IonRail applies torque on the spinning vehicle while providing thrust in a
constant direction. As we previously discussed, the turning law requires a torque from the
IonRail slider system to turn the vehicle in flight. With optimal control and carefully determining
the orbit path, the MPD thrusters slide along the tethers performing the vehicles primary
maneuvering.
-6 -4 -2 0 2 4 6
-5
-4
-3
-2
-1
0
1
2
3
4
5
Max Burn Arc With Variable Turn Rate
IonRail X Position (m)
IonR
ail
Y P
ositio
n (
m)
MPD Thrusters
Max Radius
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 240
Author: Brendon White
5.4.7 Human Factors Systems and Habitability Considerations
Scope
Incorporating humans into any space-related mission adds a great deal of complexity to every
aspect of the design, implementation and execution of the mission. These complications include
mission length, food capabilities, mass, power and volume capabilities, and psychological
aspects. Despite all of the risks associated with sending humans as opposed to robots, we still
deem it vital to use humans, as they will always play a significant role in the most remarkable
discoveries for mankind.
Background and Assumptions
Some basic assumptions were made during our design; we list some of these assumptions in
the table below:
Table 5.4.7-5 Human factors assumptions for the Crew Transfer Vehicle
Destination Ceres
Mission Length ~4 yrs
Transfer Length ~1.4 yrs
Task/Objectives
Space science/discovery
Size of Crew
Six (three male, three
female)
Safety criterion Survival of all crew members
Rather than using robotics, humans play an interactive role in this mission for several reasons:
Autonomous systems are not as reliable – since a majority of the activities that will be
performed are simply not predictable or automatic, robots would not be a sufficient choice.
Robotics are not, by nature of current human design, as productive – human beings have
creative, adaptive and cognitive brains, allowing fast decisions to be made on the spot, without
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 241
Author: Brendon White
having to wait for a specific command from elsewhere. In other words, we don’t have to be told
to do something like a robot does, we just know.
Autonomous systems and robotics are expensive and heavy - with current technology, the cost
and mass of sending the type of robotics we would need would be overwhelming [1].
Risk
Sending humans into space exploration missions is by no means easy. There are several risks
associated with doing this. The main risks that have already been addressed relate to the cost,
programmatic and biomedical categories. We focus on biomedical risks, which relate specifically
to the loss of crew safety, health and/or performance. Below is a table that summarizes some of
the risks that we identify and address in our design:
Table 5.4.7-6 Risks of incorporating humans in spaceflight missions
Severity Specification Associated Risk
Moderately
Severe
Radiation Health effects due to radiation from any source (solar flares, nuclear reactors,
etc.)
Medical Refers to any medical issue that could effect crew performance (example: a
cold)
Human Psyche Performance failure due to psychological breakdown
Bone Loss Low gravity acceleration of bone deterioration
Extremely
Severe
Environmental
Allergies, failure of artificial atmosphere, failure of purified/sterile water
Medical Toxic medication affects, space related (unexpected) antibiotic reactions
Muscle
Atrophy To the point of being unable to perform basic survival tasks or mission tasks
Sleep Performance drop due to lack of sufficient sleep
Injury Bone fractures and limited healing of such
Food Malnutrition and potentially contaminated food
Life support Supply loss, pressure loss to vessel, and/or any disruption in the artificial
atmosphere
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 242
Author: Brendon White
References
[1] Allen, C.S., Rajulu, S., Burnett, R., Cucinotta, F., Goodman, J.R. and Perchonok, M.,
―Guidelines and Capabilities for Designing Human Missions,‖ NASA Exploration Team
Human Subsystem Working Group, AIAA, 2002, pp.1-2.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 243
Author: Ben Stirgwolt
5.4.7.1 Air System
5.4.7.1.1 Air Ventilation & Circulation
Upon filtering and scrubbing the air, there needs to be a method to get the fresh air to the crew
members and to remove the ―old‖ air to back to the air filtration and scrubbing system.
Historically there are several requirements for the air circulation, which include maintaining the
airflow at a rate of 0.42 – 5.10 m3/min, keeping the exhaust velocity below 76 m/min, and
ensuring that there is sufficient fresh air at each crew member’s head [1]. In order to meet these
requirements, while maintaining a low acoustic exposure, there is a common cabin air assembly
that moves the air into central area of the CTV, and then there are small, quiet fans that move the
air into each crew member’s bedroom. There is also another small fan at the back of each
bedroom that pulls the air to the back of the room and also ensures that that there is sufficient
circulation. A possible layout of the ducting and the position of the common cabin air assembly
and the individual bedroom fans is presented in Fig. 5.4.7.1.1-1.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 244
Author: Ben Stirgwolt
Figure 5.4.7.1.1-1 The air ducting and fans in the CTV ensure that sufficient air flows to
the crew members while at the same time maintaining a low acoustic exposure.
In Fig. 5.4.7.1.1-1, the fresh air ducts are represented by the blue rectangles and the blue lines
represent the airflow. Once the fresh air leaves the common assembly in the middle of the
figure, it is pulled into each room by fans, which are represented by the green hourglass shapes.
At the back of each crew quarter, another fan pulls the air into the return ducts, which are colored
orange in the figure. At this point, the air goes through the filtration and scrubbing process and
is then re-circulated through the CTV. The mass of the ducting and fans totals 405 kg, the power
required is1.51 kW, and the ducting requires 1.75 m3.
CAD by: Brendon White, Circulation by: Ben Stirgwolt
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 245
Author: Ben Stirgwolt
References
[1] Broyan, J.L., Welsh, D.A., and Cady, S.M., ―International Space Station Crew Quarters
Ventilation and Acoustic Design Implementation,‖ 40th
International Conference on
Environmental Systems, AIAA, 2010, pp.1-2.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 246
Author: Brendon White
5.4.7.1.2 Air Tanks
Before the number and size of tanks are chosen, we calculate how much air will be needed by
the crew. The amount of oxygen needed by the crew over a given duration of time is computed
in part of our human factors program. This amount takes into account the number of crew
members, the amount of air consumed during exercise and different breathing rates for males and
females.
Our recycling system essentially has three main stages. The first stage is known as the
Sabatier stage, named after a French scientist who discovered a way to efficiently combine CO2
and H2 to form methane and water. The second stage uses the methane in a method called
Pyrolyzation. This essentially adds heat to the methane in order to produce Carbon and diatomic
hydrogen. The hydrogen is sent back into the first stage (using the Sabatier reactor) and the
Carbon is used for other applications in the mission. The final stage of the air recycling system is
probably the most well known, Electrolysis. Electrolysis takes the water created from stage two
and breaks it apart into diatomic hydrogen and oxygen. The hydrogen is sent back to the Sabatier
reactions (similar to that created in Pyrolyzation) and the oxygen is sent into the artificial
atmosphere. As the humans convert the oxygen to carbon dioxide, it is sent back into the Sabatier
reactor, combined with additional hydrogen, and the process repeats for as long as the hydrogen
supply lasts.
Below is a table giving the finalized values for how much oxygen and hydrogen we need for
the CTV:
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 247
Author: Brendon White
Table 5.4.7.1.2-1 Gases needed for artificial atmosphere
Mass, kg Tanks, kg Volume, m3
Oxygen 3507 1277 2.310E+03
Hydrogen 324.1 -- 26.69
The tanks will have a total mass of 1277 kg and will be stored in the attic, directly above the
crew living quarters. There will be a total of 10 oxygen tanks, aligned on the outside wall of the
attic. Each of these tanks has a volume of 2.147 m3, adding up to a total tank volume of 21.47
m3.
The nitrogen needed (just enough to dilute the air, making it similar to our atmosphere of 78%
nitrogen and 21% oxygen) uses one of these tanks, having a total volume of 2.147 m3. Below is a
model of the tanks by themselves and also inside of the CTV attic.
Figure 5.4.7.1.2-1 Oxygen tank used for life support systems of CTV
By: Brendon White
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 248
Author: Brendon White
Figure 5.4.7.1.2-2 Layout of the CTV attic showing the placement of the oxygen tanks along
the outer wall
By: Brendon White
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 249
Author: Jill Roberts
5.4.7.2 Water Recycling
A six-member crew needs 100,520 kg of water for drinking, clothes laundering, and hygiene
for 1.5 years. Consequently, the Crew Transfer Vehicle (CTV) would have to allow 100.52 m3
for water for each leg of the mission. This water mass is an unreasonably large cost, and could
jeopardize the feasibility of the mission. Instead of storing all this water on the CTV, we reclaim
water from multiple sources and conserve the water supply with a regeneration system. Please
see the Appendix for a demonstration on how the calculations were made and a trade study on
several water recycling system options.
The water recycling system is a hybrid between Vapor Compression Distillation (VCD) and
Multi-Filtration (MF). We show a diagram of the system in Fig 5.4.7.2 -1. Water recovered
from the crew’s urine begins the purification process with an acid pretreatment to keep the liquid
from decomposing into ammonia. It enters the VCD unit, where each of the three components
in the VCD unit rotates to provide phase separation, even in weightlessness [1].
The evaporator component of the VCD process separates the solids from the urine with heat.
We can recover 96% of the water contained in raw urine, concentrating the urine to 50% solids.
The leftover precipitate dumps into the waste water tank for processing into hygiene water in a
parallel process, described later [1].
The condenser and the condensate collector components cool and collect the separated liquid.
The water continues through the MF unit, where activated charcoal further purifies the water of
any toxins. A final post-treatment controls microbial proliferation and boosts the clean water to
drinking quality [2].
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 250
Author: Jill Roberts
Figure 5.4.7.2-1 The VCD and MF hybrid water recycling system filters water from various
sources to generate clean drinking and hygiene water.
Precipitate solids, used hygiene water, waste solids, and water vapor from the atmosphere
collect into the waste water tank for processing into clean hygiene water. It enters the MF unit,
which contains stacks of activated charcoal and filters. The filters trap solids while water passes
through the charcoal pores, purifying the water of any toxins present in the water. The water is
chemically treated to prevent microbial growth while in storage tanks [1].
The mass, power, and volume of the water recycling units depend on the number of crew
members depending on the water source. The amount of water carried onboard is dependent on
the mission time. On the Crew Transfer Vehicle, we have a crew of six people living onboard
with a particular water supply lasting 1.5 years, the duration of one mission leg. The
specifications of the water and water recycling units break down in the table below.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 251
Author: Jill Roberts
Table 5.4.7.2-1: This table shows the breakdown of mass, power, and volume of the water
supply and regeneration system.
Mass, kg Power, kW Volume , m3
Water Recycling Units 255 0.2250 0.660
Water Supply 13775 0 13.89
TOTAL 14030 0.2250 14.55
The total sum of mass, power, and volume are the contributions for the water supply and
recycling system onboard the CTV. Given that the mission out to Ceres will take less than 1.5
years, the crew has a generous supply of water until the supply can be replenished at the In Situ
Propellant Production (ISPP) stations. Please see the Appendix for a step-by-step guide of the
calculations involved.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 252
Author: Jill Roberts
References
[1] Larson, Wiley J., and Pranke, Linda K. Human Spaceflight: Mission Analysis and
Design, McGraw-Hill Companies, Inc. New York, pp. 459, 547-549, 556-559.
[2] Liskowsky, David R. Human Integration Design Handbook (HIDH), National Aeronautics
and Space Administration, Washington DC. 2010.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 253
Author: Zachary Richardson
Co-Author: Ben Stirgwolt
5.4.7.3 Food
The six person crew requires ample food in order to maintain a healthy lifestyle throughout
the journey. We have the food available in three different sources, frozen, dehydrated, and
hydroponically grown. We discuss hydroponics in the next section. The other two methods,
frozen food and dehydrated food, make up the majority of the food stock for the mission. The six
member crew requires 3.702 kg per day (dry basis) so the total mission requires a total of 5553
kg of food [1]. The crew only has enough food onboard the Crew Transfer Vehicle for the
outbound journey, and the Supply Transfer Vehicle carries the food required for the stay on
Ceres and the return trip to Earth.
5.4.7.3.1 Frozen Food Storage
We have the frozen food for the crew stored onboard the CTV in large freezers above the
crew habitat. The three freezers have a mass of 1200 kg and consume 2 kW. Due to length of the
mission, the freezers are only in use while the crew is en route to Ceres. The frozen food will
spoil if we try to have the entire mission supplied in this manner. The remaining food supply is
from either dehydrated food or from the hydroponics bay. The frozen food’s advantage over the
dehydrated is that it requires water storage and the food maintains its flavor which is a great
boost psychologically for a long term mission. The key drawback is that the frozen food requires
refrigeration which in turn requires heavy refrigerators. The initial leg of the trip includes 700 kg
of frozen food and 1633 kg of dehydrated food. This compromise minimizes the mass of the
food (from freezers) and gives a desired psychological boost for the crew.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 254
Author: Zachary Richardson
Co-Author: Ben Stirgwolt
5.4.7.3.2Dehydrated Food
We want to minimize the mass for the CTV so the food supply for the stay on Ceres and the
return trip are travelling ahead sent ahead with the STV. The frozen food requires an extra mass
for the freezers and spoils before the end of the mission. Dehydrated food replaces the frozen
food supply aboard the STV. Dehydrated food naturally requires an extra amount of water to be
produced so the ISPP stations produce the extra water required and store it on the surface of
Ceres. The total amount of dehydrated food needed for each leg of the mission can be found in
Table 5.4.7.3.1-2:
Table 5.4.7.3.1-2 Mass of dehydrated food and water required
Food, kg Water, kg
Outbound 1633 3103
On Ceres 1376 2614
Return 1923 3654
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 255
Author: Zachary Richardson
Co-Author: Ben Stirgwolt
References
[1] Hanford, A., ―Advanced Life Support Baseline Values and Assumptions Document,‖
NASA/CR—2004—208941, Aug. 2004, p. 22
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 256
Author: Ben Stirgwolt
5.4.7.4 Hydroponics
We allot 20 m3 for the production of a variety of crops. While growing crops on the CTV
causes additional mass, power, and volume as opposed to simply bring all dehydrated food, the
nutritional and psychological benefits of having fresh food during the mission cannot be
overstated. Regarding lunar base life support systems, author Peter Echart states, ―Only plants
can provide most, if not all, of the major food needs of man: calories, proteins, fats,
carbohydrates, minerals, vitamins and trace elements. It is possible for an adult person to obtain
sufficient energy on a strict vegetarian diet [1]. It is essential that on such a long mission as we
have planned, that the astronauts are reasonably comfortable, which includes what they eat. In
addition to providing fresh food, the system provides an additional daily activity for the
astronauts, perhaps breaking up the monotony of day-to-day life in the CTV. Growing plants in
a microgravity environment provides valuable information about how to best grow food for
future missions and aids the environmental control system by helping to scrub the carbon
dioxide.
The plants in the CTV grow through a technique called hydroponics. This method does not
require soil, which would cumbersome and could perhaps cause particulate build-up in the filters
because of the low-gravity environment. The hydroponics systems also can be automated to a
certain extent so that astronauts do not have to consume all of their time working to maintain the
plants. In the allotted volume, there are multiple racks, and each of the three racks have multiple
shelves which are spaced based on the height of the plants. The layout of the hydroponics bay is
shown in Fig. 5.4.7.4-1. The spacing between the racks permits the astronauts to freely move
and tend to the crops. While not shown in the diagram, there are also strawberries that are grown
from racks which are attached to the ceiling.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 257
Author: Ben Stirgwolt
Figure 5.4.7.4-1 The volume allotted for the growth of plants is filled to capacity, while
maintaining enough room for the astronauts to maneuver.
In selecting which vegetable would be grown on the CTV, we choose vegetables based on
their nutritional content and the variety of taste. This is based on the recommendation of plant
physiologist Dr. Cary Mitchell, who researches advanced life support systems for space
applications at Purdue University [2]. While there are certain crops that grow quickly, they may
not be as nutritionally beneficial, so those are foregone. The crops that grown in the hydroponics
bay include carrots, tomatoes, sweet potatoes, radishes, green onions, sweet potatoes, chard, and
strawberries.
In the hydroponics systems, the plants are not only provided with their optimal amount of
nutrients and water, but the hydroponics bay is also a closed system so that the temperature and
humidity in the growing area is monitored and maintained at prime conditions. The lighting
provided to the crops is also maintained at a level so as to provide for the best possible growing
Rack 1: Carrot, chard, tomato
Rack 2: Sweet potato, radish, green onion
Rack 3: Strawberry
By: Ben Stirgwolt
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 258
Author: Ben Stirgwolt
conditions. The hydroponics system uses LED lights that occasionally flash on and off. In the
areas where the light is reflected back, sensors then know that plants are present in that area. If
no light reflects back to the sensors, then the sensors know that there are no plants in the area,
thus the lights are turned off in this area. The intensity of the light is also changed based on the
height of the plants. Figure 5.4.7.4-2 shows of this array of LED lights and sensors in a
controlled environment.
Figure 5.4.7.4-2 LEDs and sensors determine where the light needs to be concentrated,
and the other areas are dimmed, thus saving electricity.
Even by maintaining these conditions, the hydroponics system produces only 5% of the daily
total food requirements for the crew members. However, we still believe the hydroponics system
is worth the additional mass and volume that it requires because of the aforementioned benefits.
By: Ben Stirgwolt
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 259
Author: Ben Stirgwolt
Table 5.4.7.4-1 shows the mass costs of having a hydroponics bay as opposed to bringing all
dehydrated food for the outbound trip.
Table 5.4.7.4-1 Mass comparison of food system of CTV
Biomass System , kg No Biomass System , kg
Food from Earth 2029 2333
Packaging 304 350
Water 6177 5135
Nutrients 1720 0
Structure 1591 850
Total 11821 8668
The hydroponics bay continues to produce daily crops while the astronauts explore Ceres as
well as on the return journey to Earth. During the approximately 1500 day mission, the
hydroponics bay produces 280 kg of crops on a dry basis.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 260
Author: Ben Stirgwolt
References
[1] Eckart, P., ―Science at a Lunar Base,‖ The Lunar Base Handbook, McGraw-Hill, 1999, p.
409
[2] Mitchell, Cary A. Personal Interview. 27 January 2011
Detailed Vehicle Description Crew Transfer Vehicle Page 261
Author: Brendon White
5.4.7.5 Human Required Supplies, Appliances and Medical
The biggest consideration we face in configuring the CTV layout and design is volume
allocation. To make this a human-rated mission, we must include many operational supplies,
equipment and appliances. Our spreadsheet in the CTV appendix maps out the mass, power and
volume values for each life support system on the CTV.
Table 5.4.7.5-1 CTV Crew quarters with required supplies, appliances and medical
facilities
Mass , kg Volume , m3
Food from Earth 2029 2333
Packaging 304 350
Water 6177 5135
Nutrients 1720 0
Structure 1591 850
Total 11821 8668
The crew cabin is sized to account for the volumes of appliances, supplies etc. as well as a
need for space and privacy among the crew members. We show a model of our crew cabin
design in the following figure.
Detailed Vehicle Description Crew Transfer Vehicle Page 262
Author: Brendon White
Figure 5.4.7.5-1 Model of crew habitat of the CTV
As seen in the above figure, the CTV crew living area uses a two-story configuration and can
fit inside of an Ares V payload shroud. Each floor of the crew living quarters has a volume of
138 cubic meters.
By: Brendon White
Detailed Vehicle Description Crew Transfer Vehicle Page 263
Author: Jill Roberts
5.4.7.6 Illumination
Visual lighting in the spacecraft is important to ease eye strain during work and to help
regulate the circadian rhythm in an environment without a diurnal cycle. We propose to use
solid state light modules (SSLMs), similar to one shown in the figure below, to solve this
lighting issue [1].
Figure 5.4.7.6 The SSLM solution to all our lighting problems. Drawing based on photo
from NASA specification sheet by Daniel Shultz [2].
SSLMs are composed of light emitting diodes (LEDs) and last 50,000 hours [3]. They can be
dimmed from 0-100% to accommodate multiple purposes [2]. In addition, they have a flexible
spectral power distribution. During off-duty status, astronauts can emphasize the yellow-red
spectrum to aid in falling asleep. In the morning or during duty, astronauts can increase power to
the blue spectrum to promote alertness [4]. Please see the Appendix for graphical representation
of the spectral power distribution of the lamps.
Table 5.4.7.6-1 Mass, power, and volume required to meet standards for a lighting system
aboard the CTV
Mass, kg Power, kW Volume, m3
Crew Transfer Vehicle 334.8 2.790 9.30
Drawing:Jillian Roberts
Photo: Daniel Shultz
Detailed Vehicle Description Crew Transfer Vehicle Page 264
Author: Jill Roberts
Each room in the Crew Transfer Vehicle (CTV) should meet the recommended luminous flux,
depending on the room’s purpose. The medical wing should meet the brightness level of 1000
lux [1]. The rest of the CTV can be allowed to meet requirements of standard office lighting at
500 lux. Each SSLM provides 479 lumens, has a mass of 3.6 kg, and a volume of approximately
0.1 m3. Given the lighting requirements of each room and the floor area to light, the total
number of SSLMs needed in the CTV is 93units. Twenty-one of these units are in the medical
wing alone. The total mass of the lighting system in the CTV is 334.8 kg, the volume is 9.30 m3,
and the total power required is 2.790 kW. See the above table for totals.
Detailed Vehicle Description Crew Transfer Vehicle Page 265
Author: Jill Roberts
References
[1] Hanford, Anthony J. ―Advanced Life Support Baseline Values and Assumptions
Document. ‖ Houston: Lockheed Martin Space Operations. p. 149
[2] Shultz, Daniel C. ―Solid-State Lighting Module (SSLM)‖ Kennedy Space Center. March
2008.
[3] Leveton, Lauren; Brainard, George; Whitmire, Alexandra; Kubey, Alan; Maida, Jim;
Bowen, Charles; Johnston, Smith. ―An Integrated, Evidence-Based Approach to
Transitioning to Operations: Specifications for Future Replacement Lights on ISS‖.
August 2010.
[4] Rea, Mark S.; Figueiro, Mariana G.; Bierman, Andrew; Bullough, John D. ―Circadian
Light.‖ Journal of Circadian Rhythms 8:2 (2010).
Detailed Vehicle Description Crew Transfer Vehicle Page 266
Author: Jill Roberts
5.4.7.1 Noise Suppression
Anything which creates vibrations causes noise on a spacecraft. Fans in ventilation systems
and hardware cooling are the largest source of noise. Other sources of noise include vibrations
from engines firing, fellow astronauts talking or doing work, and equipment or components
moving inside or outside the CTV. The acoustics reverberate through the ductwork and walls,
assaulting the astronauts with a barrage of noise. Over short periods of time, the excessive noise
can cause temporary hearing damage and mild physiological fatigue and stress. However, the
astronauts in our mission will be exposed to this noise over a period of several years, putting
them at risk of permanent hearing damage and severe physiological stress.
In order to reduce these effects, measures are taken to cut the amount of noise produced and
suppress the produced noise to tolerable levels. Because the ventilation system is the largest
source of long-term noise production, Ben Stirgwolt proposes a ventilation system which uses
special quiet fan. However, other design features must address vibrations from other sources
aboard the spacecraft.
Therefore, acoustic blankets are incorporated into the noise suppression design
considerations. Acoustic blankets are constructed of multi-layer materials that are sewn together
and quilted to prevent billowing. They are designed to balance acoustic abatement,
flammability, and durability. Lining the ceilings, all bedroom walls and doors, and all walls of
the common areas, the acoustic blankets significantly reduce noise reverberation. Depending on
frequency, noise levels can be reduced by 20dB or more. Because blankets also line the ceiling,
noise reverberating through the ventilation system ductwork is reduced significantly. Figure
5.4.7-1shows the locations of the acoustic blankets by the blue shading.
Detailed Vehicle Description Crew Transfer Vehicle Page 267
Author: Jill Roberts
Figure 5.4.7-1 Acoustic blankets in CTV, shown by the blue shading.
The number of layers and layer material depend on the location of the blanket. For interior
walls, including separator walls between bedrooms, the blankets consist of white Nomex ®,
Kevlar felt, and white Gore-Tex ® layers. For walls near the aisles, including the doors, the
blankets consist of white Nomex ®, BISCO ® (barium-impregnated silicon dioxide), durette felt,
and another BISCO ® layer followed by white Gore-Tex ®. Figure 5.4.7-2 pictorially shows the
blanket layers.
Detailed Vehicle Description Crew Transfer Vehicle Page 268
Author: Jill Roberts
Figure 5.4.7-2 Acoustic blanket layers depend on location in CTV
The advantage of two different acoustic blanket styles is the ability to block different noise
levels and frequencies while minimizing weight. The blankets used near aisles can block a
variety of frequencies and is especially useful for reducing noise from hallways. A layer of
BISCO ® itself provides a minimum of 11 dB reduction. The blankets in interior walls are
required to abate less noise, and thus can sufficiently reduce noise with fewer layers.
Assuming a layer of acoustic blankets on both sides of the separator walls between bedrooms,
on all interior walls of the common areas, ceilings, and near hallways, the total mass contribution
is 249 kg, with a volume of 1.67 m3. There is no power required to operate the blankets.
Detailed Vehicle Description Crew Transfer Vehicle Page 269
Author: Jill Roberts
References
[1] Broyan Jr., James L.; David Welsh, Scott M. Cady. ―International Space Station Crew
Quarters Ventilation and Acoustic Design Implementation‖. 2010. AIAA 40th International
Conference on Environmental Systems.
Detailed Vehicle Description Crew Transfer Vehicle Page 270
Author: Jill Roberts
5.4.7.8 Fire Detection and Suppression
The possibility of a fire on board the CTV requires a detection and suppression system to
catch a fire as soon as it starts, reducing damage and increasing crew safety. Smoke detectors
are placed in the intake ducts of the ventilation system [1]. The photoelectric smoke detectors
work on the principle that smoke particles scatter light [2]. A laser light beam reflects off
mirrors to photodiodes, where light obscuration and scattering measurements are taken. If the
voltage from the scattering photodiode reaches a certain level, alarms sound to alert the presence
of smoke. The crew can take action to identify the source of the fire and suppress it [1].
The portable carbon dioxide fire extinguisher (PFE) contains 2.72 kg CO2 at 5860 kPa and
discharges its contents in 45 seconds. When discharged, the bare tank and nozzle can reach
temperatures well below freezing, putting the operator’s hands at risk of damage. The
extinguisher tank is tightly enveloped by an insulating Nomex cover to keep it within tolerable
temperature limits [1].
Figure 5.4.7.8 The PFE has capabilities to extinguish many types of fires. Figure based on
photo from Alana Whitaker [1].
By: Jillian Roberts
Detailed Vehicle Description Crew Transfer Vehicle Page 271
Author: Jill Roberts
The PFE is versatile for various fire locations. It has a conical nozzle for fire suppression in
open areas. It also has a cylindrical nozzle for fire suppression ports in closed volumes. When a
fire is located in a closed area, such as behind instrument panels or between bulkheads where
electrical wiring runs, the cylindrical nozzle delivers CO2 through a small port, or hole [1].
One advantage of CO2 PFEs is that the extinguisher doesn’t need to be aimed directly at the
fire to be effective. The carbon dioxide gas arrests the reaction and suppresses the fire. In
addition, there are no particles or liquids to clean up, which could damage the electrical
equipment. The major disadvantage is that the crew is required to wear a portable breathing
apparatus (PBA) when using the extinguisher and until the atmosphere has returned to acceptable
oxygen levels [1].
Table 5.4.7.8-1 The table shows the mass, power, and volume required for a fire detection
and suppression system aboard the CTV.
Mass, kg Power, kW Volume, m3
Crew Transfer Vehicle 210.9 0.016 0.46
Assuming 11 extinguishers and 11 smoke detectors (one for each room) and 7 PBAs are
aboard the CTV, the total fire detection and suppression mass is 210.9 kg, volume is 0.46 m^3,
and 16.3 W to run the smoke detectors. The table above shows these values.
Detailed Vehicle Description Crew Transfer Vehicle Page 272
Author: Jill Roberts
References
[1] Whitaker, Alana.―Overview of ISS US Fire Detection and Suppression System.‖ NASA
Johnson Space Center 2001.
[2] Collins, Michelle M. ―Fire Protection in Manned Missions: Current and Planned.‖ Halon
Options Technical Workign Conference. Apr 2001.
Detailed Vehicle Description Crew Transfer Vehicle Page 273
Author: Jill Roberts
5.4.7.9 Waste Generation and Disposal
During the mission, the crew will go through most of the consumables on board, generating a
significant amount of waste. After recovering water waste from the waste, the remaining waste
is dead weight. By dumping unusable waste, we lose mass in the CTV, which helps us use less
propellant for maneuvers.
Sources of waste include human waste, consisting of sweat, soap and wash solids, and
biowaste [1]. See Appendix for complete breakdown of waste sources. The total waste
generated by each crew member during one day is 1.877 kg. Thus, our 6 member crew will
produce about 8219 kg of waste during a 2 year time period. This calculation assumes a plant
growing volume of 20 m3 and that food packaging is approximately 15% of the food mass [2].
We also assume the ability to recover 90% of water from the waste products, so the waste is
mostly dry solids. After water recovery, the only power required (if any) is to open the hatch to
dump waste.
Detailed Vehicle Description Crew Transfer Vehicle Page 274
Author: Jill Roberts
References
[1] NASA Exploration Team, Human Subsystem Working Group.―Guidelines and Capabilities
for Designing Human Missions‖. March 2002
[2] Hanford, Anthony J. Ph.D. ―Advanced Life Support Baseline Values and Assumptions
Document.‖ Houston, August 2004.
Detailed Vehicle Description Crew Transfer Vehicle Page 275
Author: Jillian Roberts
5.4.7.10 G-Forces for Delta-Vs
The human body can only withstand a certain amount of force for a given period of time
without grave consequences [1]. This limitation means that each phase of our mission must meet
constraints that our humans can survive. Force tolerances can vary person-to-person, depending
on size, physical strength, and flight training [2]. We assume our astronauts are in peak physical
condition and can handle a higher range of forces.
A NASA team researched several previous test cases with human subjects. These cases
involved researchers accelerating a person in particular orientation to feel a specific G-force
level along one of the body axes. The researchers measured the time the person could function
without detrimental effects, such as loss of vision or consciousness, inability to breathe, or pain
sufficient to interfere with judgment [2]. With more excessive forces, organs can rip apart and
severe injury or death may occur. See the Appendix for more details and the figure showing
limitations for human tolerance.
We don’t want our astronauts to die during the mission, especially in such a gruesome way.
Mission design and aerodynamic trajectories were constrained to keep accelerations below the
tolerable limits. The table below shows the g-forces the astronauts experience during the
mission, courtesy of Graham Johnson.
Table 5.4.7.10 G-forces for Delta-Vs during mission, data courtesy of Graham Johnson.
Mission
Phase:
Earth Kick Outbound
to Ceres
Ceres
Landing
Ceres
Launch
Return to
Earth
Aerocapture
to Earth
Force, g 0.7876 1.463 0.0558 0.1012 1.127 9.049*
*G-force at worst case scenario with high density atmosphere.
Detailed Vehicle Description Crew Transfer Vehicle Page 276
Author: Jillian Roberts
All of these forces are well below the human tolerance levels, except the aerocapture
maneuver at Earth. The 9-g of force during aerocapture is tolerable for a certain length of time.
We must check that the astronauts are not exposed to this force longer than permissible. The
graph below shows the accelerations over time during aerocapture.
Figure 5.4.7.10-2 Capsule acceleration during aerocapture peaks at just over 9 g. Graphs
courtesy of Devon Parkos.
The figure shows that the astronauts experience more than 4-g for only about 45 seconds, with
a brief max acceleration of 9-g. This acceleration is well within tolerable limits. The astronauts
could withstand twice that time at the full 9-g force. However, because the astronauts will have
spent the entire return trajectory at Mars gravity (0.38), it is unknown how they will react if we
let the mission constraints exert higher accelerations on the astronauts.
Graphs: Devon Parkos
Fig: Jillian Roberts
Detailed Vehicle Description Crew Transfer Vehicle Page 277
Author: Jillian Roberts
References
[1] Naval Aerospace Medical Institute. U.S. Naval Flight Surgeon’s Manual. 3rd
ed.
Washington D.C. 1991. pg 2-14 to 2-23.
[2] Creer, Brent Y., Captain Harald A. Smedel, Rodney C. Wingrove.―Centrifuge Study of
Pilot Tolerance to Acceleration and the Effects of Acceleration on Pilot Performance‖.
Ames Research Center. Moffett Field. 1960.
Detailed Vehicle Description Crew Transfer Vehicle Page 278
Author: Ben Stirgwolt
5.4.7.11 Humans in Artificial Gravity
Our mission specifications state that, ―artificial gravity of 0.38 g (equivalent to the gravity on
Mars) should be provided during the transfers of the crew from Earth to Ceres and from Ceres to
Earth.‖ The governing equation for producing a Martian-like gravity environment is given by
the equation 5.4.7.11-1:
(5.4.7.11-1)
where Acent is the centripetal acceleration, Ω is the angular velocity, and R is the radius of
rotation. In designing a spacecraft, one would perhaps initially assume that the best way to
achieve the desired centripetal acceleration (artificial gravity) is by changing the radius of
rotation. Reducing the radius of rotation may be good structurally and dynamically, but a shorter
radius of rotation results in a higher angular velocity (revolutions per minute). For an unmanned
space mission, having a high angular velocity may not be an issue. However, in Project Vision,
we send astronauts on a long journey where they endure countless stresses. Having
uncomfortable or sick astronauts on the CTV due to a controllable environmental factor is not in
the best interest of the success of the mission.
One of the essential questions that govern the size and configuration of the CTV is: what is an
appropriate angular velocity for the astronauts so that they are comfortable while in transit to and
from Ceres? Most of the micro-gravity research examines the effects of reduced gravity
environments on highly trained fighter pilots over a short period of time. The astronauts selected
for Project Vision also go through rigorous training, but the difference in their case is that they
will have to endure several years of living in a reduced gravity environment. Since nothing like
Project Vision has been attempted in the past, there is no research to precisely describe how
people feel and behave in artificial gravity for extended periods. Because of this, we think it is
Detailed Vehicle Description Crew Transfer Vehicle Page 279
Author: Ben Stirgwolt
best to be very conservative with the angular velocity of the spacecraft, even if that means our
structure is more massive and our spacecraft is more difficult to control.
Historically there have been five main researchers who have studied the ―comfort zone‖ of
humans in reduced gravity environments. Each researcher presented his results in a different
manner, and each researcher came up with a slightly different level of comfort. Theodore W.
Hall compiled all of the results of the historical research in 2008 and presented his results in
graphical format [1]. Figure 5.4.7.11-1 shows this complied research along with Project Vision’s
mission requirement for artificial gravity.
Figure 5.4.7.11-1 There is a consensus that the crew members will be comfortable in an
environment where the angular velocity is at 2 RPM or below.
According to the figure, 5 out of 5 researchers agree that an angular velocity of 2 RPM and
below produces a ―comfortable‖ environment for the astronauts—this is denoted by a star in the
By: Ben Stirgwolt, based on Ref. [1]
Detailed Vehicle Description Crew Transfer Vehicle Page 280
Author: Ben Stirgwolt
figure. If the angular velocity is increased to 3 RPM, then 4 out of 5 researchers think that this is
an acceptable level for the astronauts. Three out of 5 think that the upper limit for comfort is at 4
RPM. As mentioned previously, the lower the value of the angular velocity, the greater the value
for the radius of rotation. In the case of 2 RPM, the radius of rotation is approximately 85 m,
while the radius of rotation is 38 m and 21 m for 3 and 4 RPM respectively. Clearly a trade-off
exists between the comfort of the astronauts and the size of the spacecraft. However, in order to
increase the probability of success for this mission we set the RPM to 2. Reducing the RPM
reduces the uncertainty of human adaptation to artificial gravity for an extended time, enabling
the astronauts to work and live comfortably while in transit to and from Ceres
Detailed Vehicle Description Crew Transfer Vehicle Page 281
Author: Ben Stirgwolt
References
[1] Hall, Theodore W., ―Artificial Gravity,‖ Out of this World: The New Field of Space
Architecture, edited by A. Scott Howe & Brent Sherwood, AIAA, Reston, Virginia, 2009,
pp. 134-141.
Detailed Vehicle Description Crew Transfer Vehicle Page 282
Author: Ben Stirgwolt
5.4.7.12 Radiation Dosimeters and Tolerances
Astronauts are legally classified as radiation workers, and therefore NASA must employ
standards to protect them from excessive radiation exposure. Typical OSHA guidelines are
inappropriate because they are too restrictive for spaceflight activities. Instead, supplementary
regulations are used by NASA, on the basis which it applies to only a limited population and
detailed records of exposure amounts are kept. In situations where radiation exposure is
expected, hazard assessment is required and measures must be taken to keep dosage As Low As
Reasonably Achievable (ALARA). In addition, man -made radiation exposure while in-flight
must comply with Code of Federal Regulations (CFR), except when the mission cannot be
otherwise accomplished [1].
Maximum exposure limits have been drawn by NASA from recommendations made in a
report from the NCRP (National Council on Radiation Protection and Measurements). The
following table shows organ-specific exposure limits. It is based on limiting short term effects
from radiation exposure. Note that the sievert (Sv) is the official SI unit for radiation [1].
Table 5.4.7.12-1 Human organs can tolerate different levels of radiation over time [1]
Exposure Interval Depth (5 cm)
(Affects Blood
forming Organs)
Eye (0.3 cm) Skin (0.01 cm)
30 days 0.25 Sv 1 Sv 1.5 Sv
Annual 0.5 Sv 2 Sv 3 Sv
Career 1-4 Sv 4 Sv 6 Sv
To make sure the astronauts are not exceeding the maximum radiation dosage, we include
dosimeters onboard to measure the radiation. Each crew member receives a passive dosimeter,
worn like a pen in a pocket, to measure the radiation dose on each person[2]. With five units
Detailed Vehicle Description Crew Transfer Vehicle Page 283
Author: Ben Stirgwolt
located in the Crew Transfer Vehicle and one unit in the Crew Capsule are passive radiation
dosimeters, which are larger units that continuously monitor the CTV radiation. The CTV also
has one Radiation Dosimeter Assembly (RDA). It contains an area passive dosimeter, a high
rate dosimeter, and two pocket dosimeters. During a contingency event, such as a solar particle
event, the measurements can be read off the unit and reported to mission control [3].
The table below shows the total mass and volume allotted for the radiation dosimeters on the
CTV. Please see the Appendix for a breakdown of mass, power, and volume for each dosimeter
unit.
Table 5.4.7.12-2 Mass, power, and volume reserved for radiation dosimeters
Mass, kg Power, kW Volume, m3
Radiation Dosimeters 0.1345 0 0.003
Detailed Vehicle Description Crew Transfer Vehicle Page 284
Author: Ben Stirgwolt
References
[1] NASA, ―Spaceflight Radiation Health Program at JSC.‖
http://sragnt.jsc.nasa.gov/Publications/TM104782/techmemo.htm. Jan 2011.
[2] Dismukes, Kim. ―Radiation Equipment.‖
http://spaceflight.nasa.gov/shuttle/reference/shutref/crew/radiation.html. 2002. Accessed
10 Feb 2011.
[3] Johnson Space Center, NASA. ―MR004S In-Flight Radiation Monitoring with Passive
Dosimeters.‖ 2001.
Detailed Vehicle Description Crew Transfer Vehicle Page 285
Author: Alex Kreul
5.4.8. Structural Systems
5.4.8.1 Chassis
The chassis is the main structure of the Crew Transfer Vehicle (CTV), the backbone holding
all systems together. The structure consists of:
Six axial members that support each of the Primary and Earth Departure Tank Systems
and their respective tank support structure
Five hoop members that join the 6 axial members together
Tank support structure
Docking structure for parking the crew capsule when not in use (includes mounting for
the CTV power source and tether system)
Figure 5.4.8.1-1 The chassis structure of the CTV. Shown are three of the axial members,
the hoop members, and the crew capsule docking structure.
By: Alex Roth
Detailed Vehicle Description Crew Transfer Vehicle Page 286
Author: Alex Kreul
We size each component for Earth gravity, Ceres gravity, launch from Earth into LEO at
accelerations of up to 6 g’s, re-entry into Earth orbit at accelerations of up to 9 g’s, and all thrust
accelerations during the mission. More on these sizing criteria can be found in section A.5.4.8 of
the appendix.
Table 5.4.8.1-1 Chassis component breakdown. Note that the geometry sized is not exactly
the same as the geometry modeled, but that the mass estimates are consistent.
Component Material Mass, kg
Axial members (6) CFRP 3044
Tank support structure (6) Aluminum 250
Hoop members (5) Aluminum 3388
Crew capsule docking
Structure Aluminum 905
Landing legs (3) CFRP, steel 288
Total Mass (kg) 7875
5.4.8.2 Personal & Living Quarters
We place the crew quarters on the crew-related side of the CTV and surround it with a
structure of several layers. First, an inner layer of carbon fiber reinforced plastic (CFRP) acts as
a pressure vessel to maintain atmospheric pressure within the quarters. Next, a layer of high-
density polyethylene (HDPE) acts as a radiation shield. A layer of aluminum also acts as
radiation shield, as well as a micrometeorite shield. Finally, a layer of multi-layer insulation
(MLI) acts as a thermal controller.
We split he living quarters into two floors, as detailed by the CTV human factors group. One
of the floors contains a central radiation bunker, which features 2 cm thick walls of aluminum
and 12 cm thick walls of HDPE. We use this 5682 kg bunker to protect the crew in the event of
a solar particle event.
Detailed Vehicle Description Crew Transfer Vehicle Page 287
Author: Alex Kreul
Table 5.4.8.2-1 Structural dimensions for the crew quarters and storage attic. While sized
separately, these two sections are manufactured and launched as one.
Crew quarters Storage attic
Inner radius, m 3.75
Inner radius, m 3.75
Height, m 4.88 Height, m 3.77
Total height, m 8.85
Table 5.4.8.2-2 Crew quarters wall component masses. The majority of the mass present is
for radiation and micrometeoroid shielding.
Material Thickness, m Mass, kg
Inner layer CFRP 0.0005 63
Radiation shielding layer HDPE 0.06 9304
Outer layer aluminum 0.02 9288
Total Thickness, m 0.0805
Total Mass, kg 18665
5.4.8.3 Storage Space
We also place the storage attic on the crew-related side of the CTV and surround it with a
structure identical to that of the living quarters. The storage attic sits directly above the crew
quarters, allowing for easy access to stored necessities.
Table 5.4.8.3-1 Storage attic wall component masses. The majority of the mass present is
for radiation and micrometeoroid shielding.
Material Thickness, m Mass, kg
Inner layer CFRP 0.0005 106
Radiation shielding layer HDPE 0.06 7787
Outer layer aluminum 0.02 7800
Total Thickness, m 0.0805
Total Mass, kg 15693
Detailed Vehicle Description Crew Transfer Vehicle Page 288
Author: Alex Kreul
5.4.8.4 Heat Shield Insulation & Structure
The Earth re-entry heat shield is mounted on the lower side of the living quarters and storage
space structure. We require basic structural components to support the heat shield, which
experiences high pressure and drag forces upon re-entry into the Earth’s atmosphere. The total
mass of this heat shield structure is 824 kg.
5.4.8.5 Tether Cables
The tether cables of the CTV serve several purposes. First, the structural portion of the tether
withstands the tension caused by artificial gravity rotation. Second, the tether houses cables to
transfer power from the power source on the counterweight side to the personal and living
quarters, storage space, and other elements of the crew-related side. Third, the tether houses heat
pipes to transfer some of the excess heat from the power source to the personal and living
quarters. Finally, these components must be flexible enough to be reeled up. The total mass of
the structural portion of the tether is 221 kg.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 289
Author: Kim Madden
5.4.9 Thermal Control Systems
A description of the thermal control system for the CTV can be found in Section 5.7.6.2. It is the
same system as the Exploration Rover’s thermal control, with slight differences in the heater.
Heater
We include a heater in the CTV to add heat to the inside of the vehicle in case it gets too
cold for the crew. We have created a simple system to accomplish this, one that is slightly
opposite of the heat removal process. The nuclear reactor that gives power to the CTV has a low
efficiency, and thus puts out a lot of heat. We run a heat pipe through the reactor to gather this
extra heat and deliver it to the CTV, similar to a car exhaust. The heat pipe runs along the tethers,
and is made of rubber wrapped in Multi Layer Insulation (MLI), to keep heat in the water. The
rubber tubes roll up on a pulley system when the tether lengths are changed without disturbing
the water flow. Ten small radiator panels inside the vehicle are able to be manually lifted to let
heat in and closed when the temperature is comfortable for the crew. These panels are near the
air blowers to provide a heating system similar to central heat. The radiators are covered in MLI
so that heat is not added to the vehicle when is it not wanted.
Results and Summary
Table 5.4.9.-1 shows a compiled chart of the mass, power, and volume requirements for the
CTV thermal control system.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 290
Author: Kim Madden
Table 5.4.9.-1 CTV thermal control system summary
Component Mass , kg Power , kW Volume, m3
MLI Covering 58.38 0 0.21
Heat Pipe 333.86 0 3.76
Radiators 4,237.62 1.72 1.51
Aluminum Plates 28.10 0 0.01
Heater 71.36 0 0.14
TOTAL 4,729.32 1.72 5.62
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 291
Author: Austin Hasse
5.4.10 Aerodynamic Systems
5.4.10.1 Crew Transfer Vehicle Ballute
Nomenclature
A = area, m2
C = stagnation point heating coefficient, kg1/2
/m
CD = coefficient of drag
mo = mass of the Crew Transfer Vehicle
Q = stagnation point heating rate, W/cm2
Rt = torus radius, m
S = surface area, m2
T = temperature, K
V = speed, m/s
β = ballistic coefficient, kg/m2
ε = material emissivity
σ = Stefan-Boltzmann constant, W/(m2K
4)
σball = ballute areal density, kg/m2
ρ = atmospheric density, kg/m3
Subscripts
ball = ballute
The Crew Transfer Vehicle (CTV) re-enters Earth’s atmosphere by means of an aerocapture
ballute. Ballutes are a cross between a balloon and a parachute. These large area drag devices
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 292
Author: Austin Hasse
allow for reduced heating on the craft as it enters the atmosphere. They also allow the craft to
enter at a higher altitude. Ballutes provide a significant advantage over strictly aeroshells or
propellant re-entry because they offer a substantial mass reduction.
We choose to employ a trailing torus ballute for the Crew Transfer Vehicle. A clamped torus
and trailing sphere ballutes have also been investigated; however the trailing torus design allows
us to avoid some negative aerodynamic effects created from the spacecraft. Figure 5.4.10.1-1
shows the CTV and the trailing torus ballute.
Figure 5.4.10.1-1 CTV trailed by a towed torus ballute
In order to determine the required size of the ballute, we investigate the heating rates the
ballute will experience in re-entry. A larger ballute will experience less heating than a smaller
one; however it will also be have a much larger mass. We first need to determine the maximum
heating rate that the ballute can experience. This value is dependent on the material properties of
By: Austin Hasse, CTV by Alex Roth
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 293
Author: Austin Hasse
the ballute. We employ a ballute made of Kapton® polyimide film developed by the DuPont
TM
Company. This material allows has a very high maximum allowable temperature and a relatively
low density. Table 5.4.10.1-1 details the material properties of Kapton® [1].
Table 5.4.10.1-1 Material properties of Kapton®
Parameter Value Units
Max Allowable Temperature 500 oC
Material Density 1420 kg/m3
Tensile Strength 231 MPa
Material Emissivity .5 --
From the material properties we calculate the maximum heating rate that the material can
achieve. Using Eq. 5.4.10.1-1 we calculate the maximum heating for Kapton® to be 2 W/cm
2
using the emissivity and max temperature of Kapton®.
T4 = Q / (2σε) (5.4.10.1-1)
We calculate the radius of the ballute which is determined by the mass of the spacecraft, in
our case the mass of the CTV. Solving Eq. 5.4.10.1-2 for the radius of the ballute torus allows us
to size the ballute [2].
β = (mo + mball) / (CDAball) = ((mo + σballSball) / (CDAball)) (5.4.10.1-2)
We assume the coefficient of drag for a toroidal ballute to be 1.37. Equation 5.4.10.1-2 then
has only the torus radius and the ballistic coefficient as variables. We pick an arbitrary ballistic
coefficient at first to calculate the toroidal radius. We then calculate the actual heating rate on the
ballute using Eq. 5.4.10.1-3.
Q = CV3
(5.4.10.1-3)
If the heating rate on the ballute is greater than the maximum allowable heat then the ballistic
coefficient is lowered, which in turn increases the radius of the torus. Once the heating
constraints are satisfied the required ballute radius is determined. We then calculate the radius of
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 294
Author: Austin Hasse
the ballute tube. The radius of the ballute tube is determined to be a quarter of radius of the torus.
This 4:1 ratio allows us to avoid the wake created by the CTV as it passes through the
atmosphere [3]. Table 5.4.10.1-2 outlines the ballistic coefficient and radii of the CTV ballute
and Fig. 5.4.10.1-2 shows the radii of the torus from a front view.
Table 5.4.10.1-2 Ballistic coefficient and ballute radii
Parameter Value Unit
Ballistic Coefficient 1.3 --
Torus Radius (Rt) 154.68 m
Tube Radius (rt) 38.67 m
Figure 5.4.10.1-2 Front view of trailing torus with toroidal ring radius and tube radius
We then calculate the mass and volume of the various ballute components for the CTV. A
more detailed description of the calculations can be found in the appendix section A.5.4.10.1.
Table 5.4.10.1-3 and Table 5.4.10.1-4 show the tabulated values for the mass and volumes for
the CTV ballute.
By: Austin Hasse
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 295
Author: Austin Hasse
Table 5.4.10.1-3 CTV ballute masses for various components
Parameter Mass, kg
Kapton Material 2348.2
Helium Gas 1,692
Gas Tank 773.41
Tethers 989.97
Total Ballute System 4,813.6
Table 5.4.10.1-4 CTV ballute volumes for various components
Parameter Volume, m3
Ballute Expanded Volume 4,568,700
Compressed Ballute 1.65
Tank .9137
Tether 0.4446
Total Packed Ballute System 3.0083
Sizing the tethers for the Crew Transfer Vehicle consists of two parts. For the first part, we
consider the tension force needed to keep the ballute attached to the CTV. The tension required
to hold the ballute is gauged using a four tether system. The tension force depends greatly on the
material we choose for the tethers. A large tensile strength allows the tethers to remain intact
under the large amounts of stress exerted from the ballute. We choose polybenzoxazole (PBO)
fibers for the ballute tether material due to its high tensile strength and relatively small density
compared to other materials. For a more in depth analysis on tether material see appendix section
A.5.4.10.1. The material characteristics for PBO are listed in Table 5.4.10.1-5 [4].
Table 5.4.10.1-5 Polybenzoxazole material properties
Parameter Value Units
Max Allowable Temperature 650 oC
Material Density 1500 kg/m3
Tensile Strength 5650 MPa
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 296
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The second part of sizing the tethers consists of determining how to overcome tether ablation
when entering the atmosphere. Once the tether size of PBO material needed to overcome the
tension from the ballute is calculated, we calculate the amount of ablative material that needs to
be added so that the tethers do not burn up upon re-entry. The tethers use AVCOAT as the
ablative material to protect the PBO tethers.
We determine the final size of the tethers using both the size needed for tension and the
thickness of ablative material. Table 5.4.10.1-7 shows the final size of the CTV tethers and Fig.
5.4.10.1-3 shows a section view of the tether radii.
Table 5.4.10.1-7 Tether for Crew Transfer Vehicle
Parameter Radius, m
Tension Radius 0.0075
Ablative Radius 0.025
Total Tether Radius 0.0325
Figure 5.4.10.1-3 Section view of the ballute tethers showing the radii of PBO and ablative
material
By: Austin Hasse
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 297
Author: Austin Hasse
The length of the tethers is a large factor in determining the mass and volume of the tether
system. In order to determine the best length of the tethers we consider what angle the tethers
need to be so that the ablation rate on the tethers stays low and the ballute is sufficiently far back
enough to avoid the negative aerodynamic effects mentioned before. We determine the angle of
the tethers to be 60o. The final tether specifications for the CTV are listed in Table 5.4.10.1-8
Table 5.4.10.1-8 CTV tether specifications
Parameter Value Units
Number 4 --
Tether Radius 0.0325 m
Tether Length 133.98 m
Tether Volume 0.4446 m3
Tether Mass 989.98 kg
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 298
Author: Austin Hasse
References
[1] DuPontTM
Company. ―DuPont Kapton FN polyimide film.‖ kapton.dupont.com. H-38479-5,
June 2010.
[2] Gates, KL and Longuski, JM. ―Aerocapture Ballutes Versus Aerocapture Tethers of
Exploration of the Solar System.‖ Journal of Spacecraft and Rockets, Vol. 47, No. 4, July-
August 2010.
[3] Clark, I and Braun, R. ―Ballute Entry Systems for Lunar Return and Low-Earth Orbit Return
Missions.‖ Journal of Spacecraft and Rockets, Vol.45, No. 3, May-June 2008.
[4] Orndoff, E. ―Development and Evaluation of Polybenzoxazole Fibrous Structures.‖ NASA
Technical Memorandum 104814, September 1995.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 299
Author: Alexander Roth
5.4.10.2 Aeroshell and Heat Shielding
As part of the Crew Transfer Vehicle’s (CTV’s) mission, after the vehicle (and the Capsule)
returns the astronauts to Earth, the CTV will stay in Low Earth Orbit (LEO) for possible reuse.
For this to happen, the CTV will need to perform two maneuvers to slow it down. These
maneuvers involve aerobraking and Ballute aerocapture. Specifically, the aerobraking maneuver
requires that the CTV have a heat shield to protect the critical structures from the extremely high
heating rates of atmospheric reentry. If the CTV did not have a heat shield, the crew cabin side
of the vehicle would be critically damaged with no way to fix it, potentially leaving it
unsustainable for reuse ever again.
The CTV needs an aeroshell because of the chosen way to slow it down into LEO. Because
we are using an aerobraking and Ballute aerocapture maneuver, the CTV will enter the Earth’s
atmosphere to an altitude of 82.98 km (less than LEO) before regaining some altitude with a
lower velocity such that the final state of the CTV is LEO.
The CTV’s heat shield is a part of the CTV’s aeroshell, which attaches to the bottom of the
crew cabin and remains unused and unneeded for the entire duration of the mission until Earth
reentry. The heat shield is spherically shaped with a flat surface on the side that attaches to the
bottom side of the crew cabin. An image of this is in Fig. 5.4.10.2-1 below.
Figure 5.4.10.2-1 Image of CTV Aeroshell with heat shield attached (rest of CTV hidden)
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 300
Author: Alexander Roth
There were several materials that were looked at for the heat shield, but for simplicity we
choose the same material as our Capsule – Avcoat 5026-39HC/G. The discussion regarding how
this decision is made in Section 5.5.7.3 and Appendix A.5.5.7.3. (The reason why this is not
discussed here is the selection of these materials are more important for the Capsule because the
capsule undergoes much more ablation than the CTV). A table of the material’s properties is
listed below in Table 5.4.10.2-1.
Table 5.4.10.2-1 Avcoat 5026-39H/CG material properties [4]
Property Value Units
Density 5290 kg/m3
Thermal Conductivity (Isotropic) 0.24 W/m-K
Specific Heat 1610 J/kg-K
Emissivity 0.67 - -
Combustion Enthalpy 2.76E7 J/kg
Heat of Vaporization 2.65E7 J/kg
Heat of Decomposition 1.16E6 J/kg
Failure Mode Char spall - -
The dimensions of the heat shield were developed from some constraint values of other
vehicle properties. Of main concern was the aeroshell’s diameter because it could not be larger
than 8.8 m, which is the exact diameter of the Ares V extended/slightly modified cargo shroud.
In addition, the CTV’s center chassis rings are 8.8 m diameter. The other main concern was the
crew cabin has to be protected, so the aeroshell had to be larger than 8.2 m diameter as well.
Then the radius of curvature was selected to be the same as what we are using for the capsule’s
heat shield, for simplicity purposes. The aeroshell is designed to provide this function with
minimum possible mass so that useful landed mass can be maximized. An image of the
dimensions is shown below in Fig. 5.4.10.2-2.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 301
Author: Alexander Roth
Figure 5.4.10.2-2 CTV’s thermal protection system and its dimensions
Figure 5.4.10.2-3 Image of CTV’s “Compact” configuration specifically showing the
aeroshell with heat shield, the Crew Cabin, and center chassis structure
8.8 m
8.2 m
0.96033 m
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 302
Author: Alexander Roth
To determine the thickness and mass of Avcoat on the heat shield, we use Eq. 5.4.10.2-1.
(5.4.10.2-1)
Figure 5.4.10.2-4 shows the three layers of the heat shield on the CTV’s aeroshell (not to
scale). The first outside layer is Avcoat, with a thickness of 1.05 cm, as shown in dark red. The
next layer is insulation, which our code does not compute, but it often is twice the ablator
thickness, so it is 2.10 cm thick, as shown in pink. Finally, the last layer is outer wall, shown in
grey. This outer wall is the main structural component of the aeroshell and is 92.883 cm thick.
Table 5.4.10.2-2 shows the calculated thickness of Avcoat that the aeroshell will minimally need
for the aero maneuvers to work correctly and safely.
Table 5.4.10.2-2 Thickness and mass of Avcoat 5026-39H/CG for CTV
Thickness (cm) Mass (kg)
1.05 327.3581
Figure 5.4.10.2-4 Layers of TPS System (Avcoat, insulation, & outer wall) applied on the
heat shield
For the CTV to be reused, the aeroshell would have to be replaced because the ablative heat
shield would be mostly burned and nonexistent. Therefore, one of the resupply missions for a
2nd
CTV use would need to include a new heat shield and then be attached onto the CTV while it
is in LEO.
1.05 cm
2.10 cm
~93.183 cm
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 303
Author: Alexander Roth
References
[1] NASA, ―NASA Exploration Systems Architecture Study – Final Report,‖ NASA-TM-
2005-214062, November 2005.
[2] Davies, C., ―Planetary Mission Entry Vehicles Quick Reference Guide, Version 3.0,‖
NASA/SP-2006-3401, ELORET Corporation.
[3] Graves Jr., R.A., and Witte, W.G., ―Flight-Test Analysis of Apollo Heat-Shield Material
Using the Pacemaker Vehicle System,‖ NASA TN D-4713, August 1968.
[4] NASA, ―TPSX Web Edition V4,‖ Material Properties Database Web Edition V4.3 [online
database], URL: http://tpsx.arc.nasa.gov [Accessed 02/17/11].
Detailed Vehicle Descriptions Crew Transfer Vehicle Page
Author: Tony D’Mello
5.4.11 Communication Systems
5.4.11.1 Internal Communication
In Table 5.4.11.1-1, we show the communication devices that will be provided in the CTV.
Each crew member is given a television in his or her bedroom. In order to have coverage
throughout the entire circular CTV, we require three antennas placed in the center of the CTV,
each rotated 60 degrees from the previous.
Table 5.4.11.1 -1 Internal communication device characteristics for CTV
6 Televisions 6 Cell Phones 3 Antenna
Mass, kg 30 3.6 5.1e-3
Power, kW 0.6 0.42 6.3e-4
Volume, m3
0.048 0.003 5.82e-7
Detailed Vehicle Descriptions Crew Transfer Vehicle Page
Author: Tony D’Mello
5.4.11.2 External Communication
We split the crew transport vehicle’s communication requirement into two main components.
The first requirement is to transmit logistical data back to Earth and communicate with the crew
capsule during rendezvous. As we assemble the CTV in LEO, it communicates with the NASA
tracking and data relay satellites in geosynchronous orbit and continues to do so until Earth
escape. We employ a directional, Ka band, parabolic dish antenna in order to complete the link
between Earth and the tracking and data relay satellites. The high gain antenna sends logistical
information such as sensor readings, medium-quality video, and control data. Anything that is
essential to the operation and completion of the CTV transmits through this link and relays to
Earth. The specifications of this system are displayed in Table 5.4.10-1. The values listed are for
operation at the furthest distance the CTV communicates with Earth before escape. Once the
crew capsule launches, the bandwidth increases to facilitate the HDTV signal. The system is
designed such that the link can handle this increase in bandwidth since the CTV will be
considerably closer to Earth. The full link budget output is found in appendix A-5.4.11.
Table 5.4.11-7 Design parameters of the near Earth communication link
Property Value
Frequency, GHz 14.5
Data Rate, Mbps 0.488
Transmitter Receiver
Power, kW 1.00 -
Mass, kg 6.31 11.2
Diameter, m 1.50 2.00
Peak Gain, dBi 44.9 47.4
Another essential link while the CTV is in orbit around Earth is the rendezvous and docking
procedure of the crew capsule. Once the crew capsule launches and begins the ascent toward the
Detailed Vehicle Descriptions Crew Transfer Vehicle Page
Author: Tony D’Mello
CTV, it begins transmitting video and data directly with the CTV which relays all the data to
Earth. The communication link between the CTV and the crew capsule uses an ultra-high
frequency, phased-array antenna. The design details of the phased-array antenna appear in
appendix A-5.4.10 and Table 5.4.10-2. This link produces enough bandwidth to transmit two
HDTV signals of the crew as they transition from the capsule to the CTV's crew cabin.
Table 5.4.11-2 Design Parameters of the CTV to Crew Capsule Link
Property Value
Frequency, GHz 1.20
Data Rate, Mbps 50
Transmitter/Receiver
Power, kW 0.20
Mass, kg 13.5
Diameter, m 0.50
Pointing Range, deg 120
Peak Gain, dBi 28.8
The second component is providing constant HDTV between the crew and mission control on
Earth for the entire duration of the mission. This is an absolute mission requirement and serves
as the main link connecting the crew to personnel on Earth. The HDTV system is used to
monitor the status of the crew and their account of the mission, document and record all mission
activities, serve as an opportunity to communicate with family and friends over the long duration
of the mission, and provide a source of entertainment. The system also has multiple channels of
data that provide each astronaut with their own link in addition to extra bandwidth for other data
requirements. Due to the high bandwidth transmitted over such a vast distance, an optical
communication system provides the necessary ability to overcome these technical hurdles. Refer
to appendix D.3.1.1 for more information on the optical telecommunications design. The CTV
has a smaller version of the optical system present on the communication satellites. Figure
5.4.11-1 shows the receiving dish as it sits atop the crew living quarters prior to full assembly.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page
Author: Tony D’Mello
Figure 5.4.11-1 This figure shows the optical communication system in configuration before
the CTV is fully assembled. The receiver is the most notable object as it sits atop the blue
crew living quarters. The transmitter is far smaller in comparison to the other components
and can hardly be seen in the figure, but it is the small telescope opposite to the receiver on
the right side on top of the crew living quarters.
By: Alexander Roth
Detailed Vehicle Descriptions Crew Transfer Vehicle Page
Author: Tony D’Mello
While the CTV transits to Ceres, it aims the optical system directly at Earth or the Earth
trailing relay satellite until the halfway point of the transfer. Once at the halfway point, the
optical system redirects to point at one of the Ceres orbiting communication satellites. When the
CTV returns to Earth, it performs the same procedure, but in reverse. By redirecting the
telescopes mid-transfer, the overall size of the system is greatly reduced. The specifications of
the optical system are presented in Table 5.4.10-3. Figure 5.4.11-2 shows the power sizing for
the system.
Table 5.4.11-3 Design parameters of the optical communication link
Property Value
Wavelength, nm 1064
Data Rate, Mbps 232
Propagation Path Length, km 2.99 x108
Volume, m3 2.54
Mass, kg 81.5
Transmitter Receiver
Power, kW 9.27 -
Diameter, m 0.40 3.00
Length, m 0.50 1.15
Detailed Vehicle Descriptions Crew Transfer Vehicle Page
Author: Tony D’Mello
Figure 5.4.11-2 The signal to noise ratio is a value to quantify how much a signal is
corrupted by noise. The power sizing is based off of plotting the signal to noise ratio for a
given design and determining the power value where it crosses the required signal to noise
ratio. The required signal to noise ratio is related to the energy per bit to noise power
spectral density ratio and the required bit-error rate for the chosen digital modulation
method.
Detailed Vehicle Descriptions Crew Transfer Vehicle Page 310
Author: Paul Frakes
5.4.12 Rendezvous with Crew Capsule
Just prior to LEO departure, the Crew Capsule docks with the CTV. The CTV remains a
passive ―target‖ object throughout this docking procedure, whereas the Capsule is the active
―chaser‖ object. See Section 5.5.9.1 for a detailed description of this procedure.
Detailed Vehicle Descriptions Page 311
Author: Evan Helmeid
5.4.13 Ceres Operations
5.4.13.1 Landing on Ceres
From low Ceres orbit (LCO), the crew transfer vehicle (CTV) performs an optimal burn
trajectory to land on the surface of Ceres at the desired location. The optimization problem
minimizes the burn time, which is directly related to the propellant cost, and uses the steering
angle as the control. We present a detailed discussion of the optimization technique in appendix
A.5.4.13.1 and F.4.2.
We propagate the solution using MATLAB, and acquire the trajectory shown in Fig. 5.4.13.1-
1.
Figure 5.4.13.1-1 The optimal trajectory used to land the CTV on the surface of Ceres
when the crew first arrives at the celestial body.
Seen in the figure, the trajectory is smooth throughout and has a feasible final steering angle,
landing on the surface nearly vertical.
In addition to the optimal descent, we provide the astronauts with enough additional
propellant to hover for 60 seconds before landing. This buffer allows the CTV to modify its final
0 5 10 15
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Trajectory of Spacecraft
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By: Evan Helmeid
Detailed Vehicle Descriptions Page 312
Author: Evan Helmeid
landing location in case they are to land in a boulder field, on a ledge, or there is some other
difficulty. Furthermore, this allows them to maneuver to a safe distance from the ISPP station, or
to move closer, as necessary. Final propellant costs are outlined in Table 5.4.13.1-1.
Table 5.4.13.1-1 Summary of the Ceres landing specifications
Specification Value Units
Wet mass 175.1 T
Dry mass 160.6 T
Propellant mass 14.48 T
Thrust – nominal 10~100 kN
Tnominal:W range 2.115~2.306 - -
Burn time 673.0+60 s
Detailed Vehicle Descriptions Page 313
Author: Frank Fortunato & Chris Luken
5.4.13.2 Ceres Surface Operations
Landing on planetary bodies with an interplanetary spacecraft is an unconventional procedure.
Due to the design constraints for a reusable spacecraft capable of producing artificial gravity and
landing on a planet, we envision and develop a completely new design. This design is required to
withstand greater structural loads than it would in space and carry out unorthodox maneuvers in
a gravity field. The CTV lands on the surface of Ceres to complete science and other engineering
objectives. With the crew habitat near the ground, crewmembers easily access the exploration
rovers directly from the habitat.
The crew living quarters and storage area act as a base for supplies and safety during the 392
day stay on the surface of Ceres. The CTV does not move while at each ISPP station. ISPP
storage tanks fill the primary propellant tanks on the vehicle so that the CTV can launch to LCO
quickly. This keeps the tanks pressurized and allows the ISPP stations readily provide rovers and
other devices with propellant, oxygen, and water for excursions.
The components and flexibility of the CTV configuration equips it with the capability of
efficient travel. Whether short surface range or interplanetary, the CTV reaches its destination
without difficulty. This multi-role vehicle serves as the ultimate back up to the rescue rover in
the event of a failure. The Crew Transfer Vehicle contains more equipment than the rescue rover.
We take into account a certain level of risk when relocating the CTV on the surface. However,
we consider the possibility of performing such a rescue with the Crew Transfer Vehicle itself.
Detailed Vehicle Descriptions Page 314
Author: Evan Helmeid
5.4.13.3 Transfer from ISPP 1 to ISPP 2
Halfway through the stay time on Ceres, the astronauts travel from In-Situ Propellant
Production facility 1 (ISPP 1) to ISPP 2; we locate the two facilities at antipodes of Ceres. This
change in location allows for more science, and the redundancy in ISPP stations allows for a
greater margin for success.
To switch to the other side of Ceres, the crew transfer vehicle (CTV) performs an optimal
launch to ultra-low Ceres orbit (ULCO) of 25 km. After approximately half of an orbit, during
which the CTV performs a 180-degree rotation, the CTV performs an optimal landing to the
location of ISPP 2. The vehicle carries enough propellant to also perform a 60-second hover to
adjust its final landing location. The CTV uses its Ceres regime engines for this maneuver.
The resultant trajectory is sketched in Fig. 5.4.13.3-1, including locations of burns. Applicable
masses and times are presented in Table 5.4.13.3-1.
Figure 5.4.13.3-1 Ascent trajectory, coasting phase with vehicle rotation, descent trajectory,
and hover used to transfer the CTV from ISPP1 to ISPP2.
By: Evan Helmeid
Detailed Vehicle Descriptions Page 315
Author: Evan Helmeid
In the figure, the numbers correspond to the following maneuvers:
1) Optimal launch trajectory
2) Circular ULCO; vehicle rotation of 180o
3) Optimal descent trajectory
4) Optional 60-second hover maneuver
Table 5.4.13.3-1 Specifications for the ISPP transfer maneuver
Specification Value Units
Wet mass 184.6 T
Dry mass 160.6 T
Propellant mass 23.98 T
Thrust – nominal 10~100 kN
Tnominal:W range 2.0~2.3 - -
Total transfer time 85.66 min
Detailed Vehicle Descriptions Page 316
Author: Evan Helmeid
5.4.13.4 Launch from Ceres
When we launch the Crew Transfer Vehicle (CTV) from the surface of Ceres to Low Ceres
Orbit (LCO) to return home, we have a lot more mass on the vehicle than landing. The additional
mass is primarily the propellant needed for the low thrust trajectory.
Due to this extra mass, we are not able to use the Ceres regime motors; they do not provide
enough thrust to lift the CTV off of the surface. However, we do not want to add an additional
set of motors for just this single maneuver. Accordingly, we use the large, high thrust kick
motors for this launch.
The high thrust kick motors provide up to 1500 kN of thrust, which gives us a thrust to weight
ratio (T:W) of over 13, which is unnecessarily high and actually requires an excessive amount of
propellant. As such, we throttle the kick motors to 450 kN, 30% of the nominal value.
We launch the loaded CTV from the surface of Ceres to LCO, allowing us to perform system
checks before performing the large kick burn to V-infinity. We summarize the launch
specifications in Table 5.4.13.4-1.
Table 5.4.13.4-1 The masses and times required to launch the CTV into LCO while heavy
with mass for the return journey
Specification Value Units
Wet mass 471.4 T
Dry mass 424.2 T
Propellant mass 46.81 T
Thrust @ 30% 450.0 kN
T30%:W range 3.5~3.9 - -
Burn time 468.3 s
Detailed Vehicle Descriptions Page 317
Author: Trieste Signorino
5.4.14. Return Trajectory
The goals for the return trajectory of the CTV include the following:
Return to Earth and meet the mission requirement of a 2 year time limit for the crew time
in transfer
Keep the V∞ at Earth arrival under 8 km/s for aerocapture and aerobraking maneuvers to
be considered safe for the crew
These goals are the main decision making tools used to select the return trajectory from Ceres
to Earth. Appendix section A.5.4.14 provides information on the detailed process used to design
and select our final return trajectory for the CTV.
The assumptions we use in the outbound trajectory are also valid for this analysis. We again
use two types of engines, the chemical and low thrust engines. These engines will stay with our
CTV for the duration of the mission. The chemical engines provide the impulsive ∆V while the
low thrust engines provide a constant thrust throughout the return home.
Starting in an LCO of 50 km, we perform an impulsive ∆V of 2.91 km/s to escape from Ceres.
The same process we use for calculating the ∆V for the outbound trajectory is used here. The
outbound equation for determining ∆V is modified to contain the gravitational constant of Ceres
and the radius of LCO, instead of Earth parameters.
(5.4.14-4)
Detailed Vehicle Descriptions Page 318
Author: Trieste Signorino
We perform a large ∆V at Ceres, since propellant is an abundant resource for us at Ceres.
Such a large burn allows the CTV to return home faster, with less propellant on board. The mass
of propellant we use during the impulsive ∆V at Ceres was calculated using the rearranged
rocket equation found in section 5.4.3.
Unlike the Earth departure ∆V, we do not lose tanks or engines after the Ceres departure burn.
The tanks holding the Ceres departure propellant are considered the primary tanks and also hold
the propellant for the low thrust portion of the mission. We throttle the low thrust MPD’s to
provide 20 N of total thrust, while thrusting in the opposite direction of the velocity vector. The
state equations of motion used to numerically propagate this portion of the trajectory are located
in Section 5.4.3, Eq. 5.4.3-7 through Eq. 5.4.3-11. The only difference for the return trip is that
the steering angle, , is now defined by Eq. 5.4.14 -5.
(5.4.14-5)
We run our low thrust engines for 1.24 years to transfer the CTV from Ceres to Earth. The
MPDs must maintain a total thrust of 20 N to keep the CTV moving quickly towards Earth, but
still moving slow enough as to not exceed the V∞ limit set by human factors. If we thrust any
more than 20 N, we exceed this limit; if we thrust less, our transfer time increases. We arrive at
Earth with a V∞ just below the limit of 8 km/s, at 7.98 km/s. From this point, we bring the CTV
and the crew safely to Earth, with a series of aerobraking and aerocaputure maneuvers. Details of
this portion of the mission are located in Section 5.4.15. A representation of the CTV return trip
is found in Fig. 5.4.14-1.
Detailed Vehicle Descriptions Page 319
Author: Trieste Signorino
Figure 5.4.14-1 The CTV transfer from Ceres to Earth
Since an impulsive ∆V is not required to capture the CTV at Earth, we significantly reduce
the amount of propellant carried for the return trip. To take into account an increase in propellant
that would be used during a burn arc, as opposed to a purely impulsive burn for our assumption,
we increase the mass of propellant used at Ceres by 15%. The masses of propellants used during
the return trip of the CTV are located in Table 5.4.14-1.
Table 5.4.14-1 Propellant masses used for each phase of the return trajectory of the CTV.
Mission Phase Mass of Propellant, kg
∆V for Ceres Departure 203,000
Low Thrust Transfer 17,000
By: Trieste Signorino
Detailed Vehicle Descriptions Page 320
Author: Trieste Signorino
All goals for the return trajectory of the CTV were accomplished. We return to Earth in 1.24
years and therefore beat the time requirement of a 2 year maximum time limit for the crew
transfer. This trajectory is by no means an optimal solution, but was chosen to provide the
minimum amount of propellant cost after examining a select number of cases which meet the V∞
limit previously mentioned. A more detailed analysis including burn arcs and non-circular and
coplanar orbits, while using optimization techniques, would lead to an optimal solution for this
transfer.
Detailed Vehicle Descriptions Page 321
Author: Devon Parkos
5.4.15. Aerodynamic Maneuvers
v∞ = Vehicle velocity relative to Earth
Δv = Velocity change
5.4.15.1. Aerocapture Maneuver
Our CTV approaches Earth with a V∞ of 7.89 km/s. We chose a perigee altitude of 82.98 km
to ensure mission success for atmospheric density variations up to 2.5 standard deviations in
either direction. This interval corresponds with a mission success rate of 98.8%. We sized our
Thermal Protection System (TPS) to withstand both the peak heating rate that occurs during a
heighted density case and the extended ablation time that happens during a lessened density case.
The large levels of atmospheric uncertainty (See Appendix B.1.1) near this altitude
exaggerate the deviation between alternative density trajectories (Fig. 5.4.15.1-1). For this
reason, we employ the use of a ballute system, which enables our CTV design to capture with a
substantially higher V∞ than available through conventional methods. The atmospheric
uncertainty prevents an accurate optimization of entry angle, and as a result, the vehicle must
choose an initial perigee altitude with enough density to ensure aerocapture, even in the case of
maximum deviation. This leads to complications with the TPS and structural integrity.
We deploy the ballute upon approach to Earth, and it is present during the entire aerocapture
maneuver. The benefit from the large increase in surface area is that the CTV is able to capture
in a higher and less dense environment, reducing overall ablation. The ballute also increases
stability by shifting the center of pressure backwards and it generates lift, allowing the CTV to
pass through the atmosphere for a longer distance, decreasing the necessary peak acceleration for
capture.
Detailed Vehicle Descriptions Page 322
Author: Devon Parkos
Figure 5.4.15.1-1 The CTV trajectory comparison for uncertainties of 2.5σ
We defined a successful capture as a pass through the atmosphere that lowers the vehicle’s
orbital energy to that of an ellipse with an apogee radius that avoids the moon’s sphere of
influence. This maximum orbit size has a period near one week. Unfortunately, this perigee
altitude will cause severe deceleration and heightened heating rates in the elevated density case.
To compensate, we determined the worst possible ablation loss for both the tethers and the heat
shield within the 2.5 standard deviation range and sized the TPS accordingly. We also imposed a
maximum allowable acceleration of approximately 9 g’s to prevent structural failure of the CTV.
An acceleration history is available in Appendix A.5.4.15.1. The analysis of ablative material
thicknesses is in Section 5.4.10.
Detailed Vehicle Descriptions Page 323
Author: Devon Parkos
5.4.15.2. Aerobraking Maneuver and Orbit Stabilization
As visible in the figure below, the aerocapture maneuver can result in a large range of
resultant elliptic orbits. To eliminate this discrepancy and place our CTV in LEO, we execute a
carefully release of the ballute tethers. The circles denote the release locations. The resultant
orbits are nearly identical and have an apogee radius at LEO.
Figure 5.4.15.2-1 A closer view of the CTV’s trajectory near Earth
Unlike typical aerobraking maneuvers, our CTV only needs to pass through the earth’s
atmosphere one additional time to sufficiently decay its orbit. This is due to the low perigee
necessary to capture and strength of the TPS, which safely allows an additional pass through the
Detailed Vehicle Descriptions Page 324
Author: Devon Parkos
denser atmosphere. The benefit of a substantially quicker maneuver time far outweighs the
slight increase in ablative material cost.
Figure 5.4.15.2-2 The loss in specific orbital energy of the CTV during the aerocapture and
aerobraking maneuvers
To calculate the timing of the ballute release, we measure the ΔV imparted by the
aerobraking with accelerometers and integration, until the orbital energy reaches the desired
value. At this moment, we initiate the release, causing our CTV to continue on orbit similar to a
Hohmann transfer, resulting in an apogee just below LEO. The final orbit raise and stabilization
engine (FORSE) then fires, raising the perigee to LEO. We stabilize the orbit by performing
small burns with the FORSE as necessary, placing our CTV in LEO. The specific orbital energy
Detailed Vehicle Descriptions Page 325
Author: Devon Parkos
during the aerobraking maneuver for the high-density case is shown above in Fig. 5.4.15.2-2.
The figure clearly indicates the decline in orbital energy due to the aerocapture pass and the
desired reduction from the aerobraking maneuver. The terminal locations for the orbits resulting
from aerobraking for each density case are shown below. There is a slight shift in the location,
but the final circular orbit in LEO will be the same.
Figure 5.14.5.2-3 The terminal locations for the aerobraking maneuver and the location of
the primary burn of the FORSE
Detailed Vehicle Descriptions Page 326
Author: Trey Fortunato, Chris Luken
5.4.16. End of Life Configuration
After the Crew Transfer Vehicle successfully captures in an Earth orbit, the vehicle
approaches a circular orbit. The CTV remains in the parking orbit until the next mission. The
lowest mass of the vehicle occurs in this configuration. Missing components on the vehicle
include propellant tanks and kick engines, all crew consumables, and the large power source is
nearing the end of its operating period. There does remain enough power and the ability to
generate power once at Earth to power attitude determination systems and signal transmissions.
Figure 5.4.16-1 shows the Crew Transfer Vehicle in Earth orbit at the end of the mission.
Figure 5.4.16-1 CTV in final configuration awaiting next mission
The attitude thrusters perform emergency maneuvers once in Earth orbit. The Final Orbit
Raise and Stabilization Engine, FORSE, provides capability to raise or lower the parking orbit. A
By: Alex Roth
Detailed Vehicle Descriptions Page 327
Author: Trey Fortunato, Chris Luken
future mission would require replenishment of consumables, new power source, and replacement
propellant tanks. Table 5.4.16-1 shows the breakdown of only replacement systems and their
masses.
Table 5.4.16-1 Recoverable and non-recoverable replaced systems
System Mass, kg Percent of Total Dry Mass, %
Primary Tanks 12,885 9.24
Earth Departure Tanks 14,318 10.26
Power Source 25,872 18.55
Crew Consumables 27,105 19.43
Crew Capsule* 9836 7.05
Total Replacement 80,180 57.48
We replace the non-recoverable propellant tanks and their attached engines for future
missions. The heat shield only protects the crew habitat and center chassis structure. The non-
recoverable replaced systems comprise 19.5% of the total vehicle dry mass. The recoverable
systems that require replenishment amount to 37.98% of the total vehicle try mass. This indicates
that a large portion of the vehicle is directly being used or consumed, resulting in a lower
effective inert mass. This design allows the mission greater chance to succeed such that we
require minimal overall mass to complete the mission.
With the possibility of future missions, we look to various other destinations for the CTV.
Different size tanks and their attached propulsion systems would provide the ability to travel to
different locations in the solar system directly from Earth. The chassis design used adjusts for
future changes and evolutions in design. The method of attaching and removing propellant tanks
carrying engines opens doors to the future of manned spaceflight to various destinations.
Detailed Vehicle Descriptions Page 328
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Detailed Vehicle Descriptions Crew Capsule Page 329
Author: Jillian Roberts
5.5 Crew Capsule
5.5.1 Crew Capsule Configuration Overview
Because of the similar size of crew, the Crew Capsule was sized using the Orion Spacecraft as
a baseline for radius of curvature and capsule angles. Given a required internal volume that the
capsule would have to hold and volumes for the ballute, tank, and parachute the capsule was
sized accordingly. First, the ballute and parachute cap is sized, accounting for a tunnel for
astronauts to climb through of width 0.75 m. Once this cap is sized, the rest of the capsule is
built next to it, making sure everything will fit in the capsule, including extra space for astronaut
movement within the capsule. The table below shows all the important dimensions of the
capsule. Please see the Appendix for a detailed drawing of the Capsule dimensions.
Table 5.5.1-1 Key dimensions for the Crew Capsule
Description Unit Dimension
Overall Length, including heat shield m 3.986
Length of capsule without heat shield m 3.425
Length of full ballute and parachute
cap
m 0.75
Length of ballute and tether storage m 0.436
Tunnel width (diameter) m 0.75
Angle of decreasing capsule radius deg 32.4
Base diameter of capsule m 5.25
Radius of curvature of heat shield m 6.202
At the time the capsule was sized, it was not known whether or not the aerodynamically
curved bottom could provide usable space. Therefore, once the heat shield was added to the
capsule sizing, we gained approximately 10 m3 of volume for contingency storage. The total
mass, power, and volume parameters of the Capsule are found in the table below.
Detailed Vehicle Descriptions Crew Capsule Page 330
Author: Jillian Roberts
Table 5.5.1-2 Mass, power, and volume for the Crew Capsule
Mass, kg Power, kW Volume, m3
Capsule 9834 1.89 43.7
The interior of the capsule is designed so that the six crew members can live relatively
comfortably for as long as a week. The seats are arranged symmetrically around the axis of the
capsule, with each astronaut seated towards the tunnel and the heat shield at their backs. The
seats are position so that the astronauts will be able to manage the high g-forces during re-entry.
In addition, there is sufficient room for storage of supplies, food, and personal items. There is
extra “moving” in the space behind the heat shield, making the capsule a little bit more spacious
in case the crew members need to live in the capsule for a week.
Detailed Vehicle Descriptions Crew Capsule Page 331
Author: Joel Lau
5.5.2. Crew Capsule Power
We equip the Crew Capsule with two Sodium-Sulfur (Na-S) batteries to provide power. The
Mass, Power and Volume of the Crew Capsule power system are tabulated in Table 5.5.2-1
Table 5.5.2-1 Two Na-S batteries provide power to the Crew Capsule
Power Provided, kW 2.000
Battery Life at Peak Power, hr 10.59
Total Energy Storage, kW-hr 20.00
Mass, kg 129.03
Volume, m3 0.2775
As shown, the maximum power provided is 2.000 kW. The majority of this powers the
instrument panel, which requires the most power, 1.500 kW. The next highest power need is the
communications hardware (antennas and transponder), which totals 310 Watts. The crew
capsule’s total power budget is 1.889 kW and the distribution is shown in Table 5.5.2-2.
Table 5.5.5-2 Crew Capsule Power Budget
System Power Requirement, W
Attitude Control Hardware 40
Transmit/Receiving Antenna 200
UHF Transponder and Duplexer 110
Bio-Suit Pressure Suits 17.6
Fire Detection and Suppression 1.5
Instrument Panel Allowance 1500
Coolant, Coolant Pump 20
Total 1889.1
For the majority of its life, the Crew Capsule is connected to the Crew Transfer Vehicle
(CTV) and it is powered by the CTV power system. During launch and re-entry, the Crew
Capsule is disconnected from the CTV for a maximum of ten hours. At this time, the Na-S
Detailed Vehicle Descriptions Crew Capsule Page 332
Author: Joel Lau
battery array powers the Crew Capsule. Upon reconnection to the CTV, the CTV’s power system
again powers the Crew Capsule and also recharges the Na-S batteries.
The power solution was designed to provide slightly more than the maximum requirement.
The Na-S battery array supplies a total of 2.000 kW of peak power, slightly in excess of the
1.889 kW required. Additionally, at peak load, the batteries have a lifespan of 10.59 hours,
longer than the 10 hours that the capsule will be disconnected from the CTV in normal
circumstances.
We also designed the system to be redundant. The array includes two batteries, so that if one
battery fails, the other battery still provides enough power for communication and for vital safety
systems. This means that even in the event of a partial power failure, the coolant system and fire
suppression system, which protect the Crew Capsule from destruction, are still active. Most
importantly, the communication system will still be powered, and the astronauts will be able to
communicate with Earth or the CTV as needed.
We chose Na-S batteries because they have several advantages over other power solutions.
The most obvious is their high energy density. The batteries aboard the crew capsule have an
energy density of 155 W-hr/kg, and a mean density of 465 kg/m3. This is more than double that
of Nickel-Metal Hydride (NiMH) batteries (70 W-h/kg) [1], the most commonly used batteries in
space applications. Na-S batteries also have a highly efficient depth of discharge, approximately
90%, compared to 80% for Ni-H batteries. Depth of discharge is a measurement of how low a
batteries charge can be before it must be recharged. Na-S batteries are at a Technology Readiness
Level of seven, a Na-S battery, designed for space applications, was tested successfully aboard
the Space Shuttle in November 1997 [2].
Detailed Vehicle Descriptions Crew Capsule Page 333
Author: Joel Lau
Resources
[1] Wertz, James Richard, and Larson, Wiley J. "Space Mission Analysis and Design",
Microcosm, 1999.
[2] “NRL NaSBE Experiment”, 1997, accessed 02 Apr 11 at
[ http://www.nrl.navy.mil/pao/pressRelease.php?Y=1997&R=82-97r]
Detailed Vehicle Descriptions Crew Capsule Page 334
Author: David Wyant
5.5.3 Propulsion Systems
An integral part of any human spacecraft is the launch abort system. To this end, the final
crew capsule design includes a launch abort system (LAS) for crew safety. We design the LAS
with the same performance parameters in mind as those used to design the Orion Crew Capsule.
Both in size and weight, this previously designed capsule serves as a useful analogue and starting
place for this analysis.
We will deploy an abort system with the same ΔV requirement of 264 m/s with a more
efficient propellant on our crew capsule [1, 2]. The LAS will also have a similar burn time (3
seconds). An HTPB propellant, solid rocket motor was used in the final design for safety
reasons. In general, maximum thrust is achieved more quickly with a solid motor than a liquid
motor. This is desirable in emergency situations as immediate response times are necessary.
The Launch Abort System was designed to fulfill the same requirements as the Orion Launch
Abort System. The ΔV requirements for the LAS were taken from those requirements and the
new abort system accordingly sized.
We size the motor case according to the method and procedures in Space Propulsion and
Design [3]. This analysis yields a system with the dimensions and characteristics shown below.
The detailed breakdown of this analysis is presented in the appendices.
Detailed Vehicle Descriptions Crew Capsule Page 335
Author: David Wyant
Figure 5.5.3 - 1 Sketch of the solid rocket abort system. The grey area is propellant and the
outline is the motor case.
Table 5.5.3-1 Engine performance and design
Mass, kg Volume, m3 Isp, s
HTPB Propellant 1031 0.600 270
Pressure Vessel 15.12 - - - -
Skirt
Motor Case
Nozzle
7.56
24.95
66.91
- -
0.6226
--
- -
- -
- -
Total System 91.87 0.6226
Detailed Vehicle Descriptions Crew Capsule Page 336
Author: David Wyant
References
[1] Wade, M., "Orion LAS," Astronautix, [http://www.astronautix.com/craft/orionlas.htm.
Accessed 3/31/11.]
[2] “Orion Launch Abort System,” Orbital,
[http://www.orbital.com/NewsInfo/Publications/Industry_Pad_Fact.pdf. Accessed 3/31/11.]
[3] Humble, R.,Henry, G., Larson, W., Space Propulsion Analysis and Design, Mc-Graw Hill,
1995.
Detailed Vehicle Descriptions Crew Capsule Page 337
Author: Jillian Roberts
5.5.4. Human Factors Systems and Habitability Considerations
Because the astronauts spend up to 10 days in the Crew Capsule, we must make water
provisions to allow for drinking, food rehydration, and basic hygiene while away from the
CTV. We conducted a trade study which determined that storing water instead of recycling it,
would significantly decrease mass. While in the capsule, the crew uses the minimum amount
of water required to reduce mass until rendezvous with the CTV. This trade study can be
found in the Appendix. The total mass, volume, and power requirements for the Crew
Capsule are found in the table below.
Table 5.7.4-1 Specifications for the water supply and recycling system
Crew
Members
Days Mass, kg Power, kW Volume, m3
Water Supply and
Regeneration
6 10 172.5 0 0.175
In case of a fire, the Crew Capsule has two fire extinguishers and one smoke detector. See the
Fire Suppression and Detection section from the Crew Transfer Vehicle and its corresponding
Appendix for details. The mass, power, and volume parameters can be found in the table below.
Table 5.4.7-2 Specifications for the fire suppression and detection system
Mass, kg Power, kW Volume, m3
Fire Detection and Suppression 23.27 0.0015 0.0788
Detailed Vehicle Descriptions Crew Capsule Page 338
Author: Jillian Roberts
To provide an ergonomic working environment which is well-lit, the Exploration Rovers will
have a lighting system. The table below describes the mass, power, and volume of the lighting
system. We assume the Crew Capsule needs 200 lux, which is recommended office lighting, for
the astronauts to efficiently perform basic tasks in the Capsule.
Table 5.4.7-3 Specifications for the lighting system
Mass, kg Power, kW Volume, m3
Lighting System 21.6 180 0.600
To make sure the astronauts aren’t getting exposed to too much radiation, there is a Passive
Radiation Dosimeter (PRD) aboard the capsule. The dosimeter’s mass, power, and volume are
described below. For addition detail, please refer to the section on Radiation Dosimeters and
Tolerances in the CTV vehicle description.
Table 5.4.7-4 Specifications for the radiation dosimetry system
Mass, kg Power, kW Volume, m3
Radiation Dosimeters 0.1345 0 0.000198
The astronauts will need food for 10 days, and the corresponding mass, power, and volume
for food supplies is in the table below. Please see the food section in the CTV vehicle
description for additional details.
Table 5.4.7-5 Specifications for the food
Mass, kg Power, kW Volume, m3
Food System 138.9 0 9.01
Detailed Vehicle Descriptions Crew Capsule Page 339
Author: Jillian Roberts
The astronauts will need air for 10 days, and the corresponding mass, power, and volume for
breathing air and tankage is in the table below. Please see the air section in the CTV vehicle
description for additional details.
Table 5.4.7-6 Specifications for the air system below
Mass, kg Power, kW Volume, m3
Air System 30.53 0 9.01
Detailed Vehicle Descriptions Crew Capsule Page 340
Author: Paul Frakes
5.5.5 Attitude Determination and Control Systems (ADCS)
We equip the Crew Capsule with an inertial Motion Reference Unit and associated computer
system, which serves as the attitude determination system. The attitude control system consists
of four attitude thrusters, each attached to Canfield joints (see Fig. 5.5.5-1).
Figure 5.5.5-1 Model of a Canfield joint. The payload of the distal plate can be maneuvered
through 2π steradians. The central propellant feed lines are flexible.
Canfield joints, described in detail in Section 5.2.6, enable us to reduce the number of
required attitude engines to four from the traditionally required 16, since they can be gimbaled to
point in any direction within a hemisphere, i.e., the range of motion is 2π steradians. The joints
are also controlled by the attitude determination computer system.
We place the engines along the center of mass of the Crew Capsule, so that side-to-side
motion of the Capsule can be accomplished without creating a pitching moment. This location
was determined to be approximately one quarter of the distance from the bottom of the Capsule
to the top, and this location lies along the axis of symmetry of the Capsule. The engines produce
By: Alex Roth and Paul Frakes
Detailed Vehicle Descriptions Crew Capsule Page 341
Author: Paul Frakes
10 Newtons of thrust each, and we employ hypergolic propellants monomethylhydrazine (MMH)
and nitrogen tetroxide (NTO). Each engine has an Isp of 220 seconds.
The mass, power, and volume requirements for the Crew Capsule ADCS is given below in
Table 5.5.5-1.
Table 5.5.5-1 Mass, power, and volume requirements for Crew Capsule ADCS
Mass, kg Power, kW Volume, m3
Propellant 88.16 0 0.074
Attitude Control
Hardware
57.78 0.04 0.95
Override Joystick
and Computer
0.0249 0.00 0.00
Total 145.96 0.04 1.024
The attitude control thrusters will be employed in each of the maneuvers required of the Crew
Capsule (see Section 5.5.9). For the majority of its life, the Crew Capsule will be attached to the
CTV, which lessens the role of ADCS onboard the Crew Capsule. Other perturbations, such as
environmental forces and torques that act on the Capsule, were not considered because the
Capsule spends so little time subjected to these forces. Also, no active attitude control systems
are required for the atmospheric entry phase, since the ballute keeps the Capsule in the proper
orientation throughout entry. Therefore, the only attitude control propellant that is needed is what
is required for the maneuvers described in Section 5.5.9.
Detailed Vehicle Descriptions Crew Capsule Page 342
Author: Andrew Curtiss
5.5.6. Structural and Thermal Systems
5.5.6.1 Structures Overview
We base the design of our Crew Capsule model on the Orion crew module, which is Project
Constellation’s main crew capsule [1]. Our capsule is larger than the Orion capsule because it is
required to support six astronauts while Orion only needs to support four people. The mission of
the Crew Capsule is to transport the astronauts from Earth to the CTV and from the CTV back to
the surface of the Earth. During the trip to Ceres, the Crew Capsule is uninhabited and remains
that way until the astronauts have returned from Ceres to low Earth orbit (LEO).
During re-entry (the trip from LEO to the surface of Earth), the capsule flies through Earth’s
atmosphere. The friction of reentry produces enough heat that, if left unprotected, the Crew
Capsule would burn up. The heat shield protects the crew from the heat of reentry and is a main
component of the structural design. The Capsule also contains a pressurized cockpit that the crew
inhabits during the Crew Capsule mission. Finally, we design the Capsule to feature a container
which holds the science payload from Ceres: one ton of regolith from the surface.
Detailed Vehicle Descriptions Crew Capsule Page 343
Author: Andrew Curtiss
5.5.6.2 Regolith Storage Container
We require a container to store the cryogenically frozen regolith for the trip from Ceres back
to Earth. We place the container inside the Crew Capsule where it reenters with the crew
members. Of the 1,000 kg of Ceres rock that we require to be returned to Earth with the crew;
500 kg comes from each of the ISPP stations. We also require that the regolith be frozen
cryogenically on the return trip so that the rock remains in the condition it is in on the surface of
Ceres. In addition, we separate the samples from each of the ISPP stations using a partition in the
storage container.
We model the container as a simple rectangular prism as shown in Fig. 5.5.6.2-1. We make
the entire container out of multilayer insulation (MLI). The insulation keeps the heat of reentry
from warming the rock. The thickness of the MLI in the walls and partition is 0.0169 m which is
roughly twice as thick as the walls of the propellant tanks.
Figure 5.5.6.2-1 Depiction of the container in which Ceres regolith is returned to Earth.
The thickness of the container’s insulation keeps the rock cold.
The figure shows the dimensions L, w, and h which refer to length, width, and height,
respectively. The dimensions satisfy the volumetric requirements imposed by the 1,000 kg of
regolith to be returned. The box is capable of holding the regolith and keeping it frozen for the
By: Andrew Curtiss
Detailed Vehicle Descriptions Crew Capsule Page 344
Author: Andrew Curtiss
flight from the CTV back to the surface of the Earth inside the Crew Capsule. We list these
dimensions, along with the thickness of the container, the mass, and the volume in Table 5.5.6.2-
1. Note that the table does not include the mass of the regolith which is returned.
Table 5.5.6.2-1 Parameters of Ceres regolith sample return box
Unit Value
Length m 1.2705
width m 1.2705
height m 0.6352
thick m 0.0169
Mass kg 180.7491
Volume m3 1.0225
Detailed Vehicle Descriptions Crew Capsule Page 345
Author: Andrew Curtiss
5.5.6.3 Structural Mass
The Crew Capsule’s structural mass features two main components: the outer shell and the
heat shield backing structure. We model the outer shell as a partial cone with a thickness of 4 cm.
The outer shell’s structure supports the pressure inside the capsule and the loads that the capsule
undergoes during flight and reentry. We design the shell out of Aluminum because of its
relatively low weight and its high strength.
The heat shield backing is the second main structural component of the Crew Capsule. We
design the backing structure to attach the heat shield to the capsule and to support the dynamic
pressure loading during reentry. We estimate the mass using purely historical data from other
reentry capsules. The mass is fully incorporated in the existing structure, so there is no volume
added for the backing structure. In addition to the structural mass, we add multi-layer insulation
to insulation protect the capsule from the radiant heat coming from the shield. The structural
mass breakdown is summarized in Table 5.5.6.3-1.
Table 5.5.6.3-1 Breakdown of structural mass components
Component Mass, kg Volume, m3
Shell 5907.5 2.188
Backing 295.82 0
Insulation 5.58 0
Detailed Vehicle Descriptions Crew Capsule Page 346
Author: Andrew Curtiss
References
[1] “Constellation Orion Crew Exploration Vehicle”, NASA Fact Sheet No. FS-2008-07-031-
GRC, January 2009.
Detailed Vehicle Descriptions Crew Capsule Page 347
Author: Kim Madden
5.5.6.4 Thermal Control System
The details for the Crew Capsule thermal control system can be found in Section 5.7.11.2.
Table 5.5.6.4-1 shows a compiled chart of the mass, power, and volume requirements for the
Crew Capsule thermal control system.
Table 5.5.6.4-1: Crew Capsule Thermal Control System Summary
Component Mass, kg Power, kW Volume, m3
Heat Pipe 6.85 0 0.98
Radiators 53.79 0 0.05
Aluminum Plates 7.03 0 0.008
TOTAL 67.67 0 1.04
Detailed Vehicle Descriptions Crew Capsule Page 348
Author: Kim Madden
References
[1] Birur, G. C., Siebes, G, and Swanson, T. D., “Spacecraft Thermal Control”, Encyclopedia
of Physical Science and Techonology, 3rd
ed., Academic Press, 30 March 2001.
[2] Holman, J., Heat Transfer, 10th
ed., McGraw-Hill, New York, 2009, Chaps 8-10.
Detailed Vehicle Descriptions Crew Capsule Page 349
Author: Austin Hasse
5.5.7 Crew Capsule Aerodynamic Systems
5.5.7.1 Re-entry Ballute
Nomenclature
A = area, m2
CD = Coefficient of drag
mo = mass of Crew Capsule
Q = stagnation point heating rate, W/cm2
S = surface area, m2
T = temperature, K
β = ballistic coefficient, kg/m2
ε = material emissivity
σ = Stefan-Boltzmann constant, W/ (m2K
4)
σball = ballute areal density, kg/m2
Subscripts
ball = ballute
We designed the Crew Capsule re-entry ballute to be a large trailing toroidal ring tethered to
the Crew Capsule. This design is the same as that of the ballute on the Crew Transfer Vehicle.
Fig. 5.5.7.1-1 gives a graphic portrayal of a trailing torus ballute attached to the Crew Capsule.
Detailed Vehicle Descriptions Crew Capsule Page 350
Author: Austin Hasse
Figure 5.5.7.1-1 Capsule ballute deployed for re-entry.
One of the limiting factors on the size of our ballute is the choice of material in which the
ballute is fabricated. We choose to use the Kapton® polyimide film developed by the DuPont
TM
Company. Kapton® has very desirable material properties for ballutes because of the maximum
allowable heat and the relatively low density. We calculate the maximum allowable heating rate
on the ballute from the material properties of Kapton®. These properties are listed in the
appendix section A.5.4.10.1-1. We employed a thickness of 7µm for the ballute material. This
extremely thin layer of Kapton® allows us to keep the mass of the ballute material low. We
established the max heating rate of 2 W/cm2 for Kapton
®.
By: Austin Hasse, Capsule by Alex Roth
Detailed Vehicle Descriptions Crew Capsule Page 351
Author: Austin Hasse
The size of the ballute is determined in the same way as the CTV ballute, which employs the
mass of the object in which it is designed and the ballistic coefficient. The ballistic coefficient
regulates how large the radius of the ballute must be to effectively slow the craft in the
atmosphere. This coefficient is determined by the heating rate on the craft. We evaluate the size
of the ballute using Eq. 5.5.7.1-2.
β = (mo + mball) / (CDAball) = ((mo + σballSball) / (CDAball) (5.5.7.1-2)
From this equation we derive the radius of the toroidal ring. The radius of the ballute tube is
found to be one forth that of radius of the torus. This configuration allows the wake of the
capsule to go through the inner torus ring without disturbing the flow around the ballute. Figure
5.5.7.1-2 depicts the radii of the torus.
Figure 5.5.7.1-2 Front view of trailing torus with toroidal ring radius and tube radius.
By: Austin Hasse
Detailed Vehicle Descriptions Crew Capsule Page 352
Author: Austin Hasse
The ballute is pressurized using helium gas. We choose a deployed pressure of 10 Pa. The
pressure of the atmosphere at our capture altitude is 1.037 Pa. The helium is stored in a tank
aboard the capsule which is pressurized to 50 MPa to reduce the size of the storage tank and is
made of titanium.
Values for the masses and volumes of the ballute structure are given in Table 5.5.7.1-2 and
Table 5.5.7.1-3 respectively.
Table 5.5.7.1-2 Mass values for various ballute components
Parameter Mass, kg
Material Mass 181.4
Tank Mass and Helium 13.54
Tethers 75.88
Total Ballute System 270.82
Table 5.5.7.1-3 Volume values for various ballute components
Parameter Volume, m3
Expanded Ballute (m3) 98,094
Compressed Ballute (m3) 0.1277
Helium Tank (m3) .0051
Tethers (m3) 0.0358
Total Packed Ballute Volume (m3) 0.1686
The sizing for the Crew Capsule tethers is done in the same manner as that of the CTV (see
report Section 5.4.10.1). The Crew Capsule ballute also employs the doubled layer tethers
consisting of a PBO layer and an ablative layer. The size of the Capsule tethers is much smaller
than the CTV tethers because the Capsule experiences less tension and heating upon re-entry.
Table 5.5.7.1-4 gives values for the radii of the tether layers.
Detailed Vehicle Descriptions Crew Capsule Page 353
Author: Austin Hasse
Table 5.5.7.1-4 Tether radii for the Crew Capsule
Parameter Radius, m
Tension Radius 0.0025
Ablative Radius 0.015
Total Tether Radius 0.0175
We calculate the length of the tethers at a 60o angle with respect to the radius of the ballute,
the same calculation used to calculate the length of the CTV tethers. Using this length, the tether
specifications for the Crew Capsule are obtained. The measurements of the Capsule tethers are
listed in Table 5.5.7.1-5.
Table 5.5.7.1-5 Crew Capsule tether specifications
Parameter Value Units
Number 4 --
Tether Radius 0.0175 m
Tether Length 37.24 m
Tether Volume 0.0385 m3
Tether Mass 75.88 kg
Detailed Vehicle Descriptions Crew Capsule Page 354
Author: Austin Hasse
5.5.7.2 Crew Capsule Parachute
Nomenclature
Aparachute = area of parachute, m2
CD = coefficient of drag
D = force of drag, N
V = speed, m/s
ρair = density of air, kg/m3
The parachute system for the capsule comprises three separate chutes, each connected to the
top of the Capsule. We determined a three chute system was ideal for the capsule re-entry due to
the added safety provided from multiple chutes. Figure 5.5.7.2-1 illustrates how the three chute
system is deployed.
Figure 5.5.7.2-1 Re-entry with deployed Crew Capsule parachute system
We determine the size of the parachutes needed to land the Capsule safely using Eq. 5.5.7.2-1.
By: Austin Hasse, Capsule by Alex Roth
Detailed Vehicle Descriptions Crew Capsule Page 355
Author: Austin Hasse
D = ½ ρairV2AparachuteCD (5.5.7.2-1)
The required drag needed to slow the capsule down is determined from the mass of the
capsule and the maximum acceleration that it experiences after the parachutes are released. We
size our parachutes such that any combinations of two chutes create enough drag for the craft to
return safely to the surface. We create an extra failsafe by employing three chutes because if one
single chute fails the other two chutes remain. Table 5.5.7.2-1 details the parachute
specifications for the two-chute system with a redundant third “safety chute.”
Table 5.5.7.2-1 Crew Capsule parachute specifications
Parameter Value Units
Parachute Radius 15.78 m
Single Chute Mass 179.91 kg
Single Chute Volume 0.1564 m3
Total Chute Mass 539.74 kg
Total Chute Volume 0.4693 m3
Detailed Vehicle Descriptions Crew Capsule Page 356
Author: Alexander Roth
5.5.7.3 Thermal Protection System (TPS)
As part of the capsule’s mission, after the Crew Transfer Vehicle (CTV) returns the astronauts
most of the distance to Earth, the astronauts will transfer to the capsule for the reentry to Earth.
They will leave the CTV in Low Earth Orbit (LEO). Therefore, once the astronauts transfer over
from the CTV, the capsule will deploy its Ballute for aerobraking and then use its Thermal
Protection System’s (TPS’s) heat shield to protect the vehicle and its contents for the actual
reentry and landing. If the capsule did not have a heat shield, the capsule would not survive
reentry.
The heat shield attaches to the bottom side of the capsule and remains unused and unneeded
for the entire duration of the mission until Earth reentry. The heat shield is spherically shaped
and fits under the astronauts’ feet, below the floor of the capsule. Figure 5.5.7.3-1 shows this.
Figure 5.5.7.3-1 Image of Capsule Heat Shield (Rest of body hidden)
The diameter was specifically constrained to be the capsule’s exact diameter because any
diameter larger would be inefficient and any diameter smaller would not protect the whole
vehicle. Diameter of ARC Capsule = 5.25 m and Fig. 5.5.7.3-2 shows this.
Detailed Vehicle Descriptions Crew Capsule Page 357
Author: Alexander Roth
Figure 5.5.7.3-2 Capsule and heat shield full assembly
The aeroshell was designed using the Apollo, Mars Exploration Rovers (MER), and Orion
capsules as references. Table 5.5.7.3-1 below shows all these diameters, radius of curvatures,
and relative vehicle mass for comparison.
Table 5.5.7.3-1 Comparison of our capsule’s dimensions with NASA’s past capsules
Vehicle Diameter , m Radius of Curvature, m Vehicle Mass, kg
Apollo 4 [2] 3.91 4.690 5,424.9
Mars Exploration Rovers (MER) [2] 2.65 N/A (70°Sphere Cone) 836.0
Orion CEV [1] 5.00 6.476 8,913.0
ARC (Our Capsule) 5.25 6.202 9,836.5
Detailed Vehicle Descriptions Crew Capsule Page 358
Author: Alexander Roth
There were several materials that were looked at for the heat shield and their material
properties are listed in Appendix A.5.5.7.3. These materials are low in weight and have a high
maximum temperature rating. The material chosen for the heat shield was Avcoat. Avcoat was
selected because of the various materials it had the lowest density with the highest isotropic
thermal conductivity, the highest specific heat, and the best emissivity value. In addition, even
though Avcoat has not been in production for 20+ years, NASA recently restarted a small
amount of production for use during testing for the ill-fated Project Constellation.
Table 5.5.7.3-2 Comparison of heat shield materials of NASA’s past capsules
Vehicle Material Thickness, cm
Apollo 4 [2] Avcoat 5026-39 HC/G 4.32
Mars Exploration Rovers (MER) [2] SLA-561V 1.57
Orion CEV [1] Avcoat - ARC (Our Capsule) Avcoat 5026-39HC/G 1.05
Avcoat is an epoxy-novolac resin with special additives in a fiberglass honeycomb matrix.
The char of the material is composed mainly of silica and carbon.[1] It is manufactured directly
onto the Capsule’s substructure. Avcoat uses ablation to absorb and dissipate the heat applied to
it during atmospheric reentry. Ablative materials are designed to burn away slowly and in a
controlled manner so that the heat is carried away from the spacecraft by the gases in the upper
atmosphere and the gases released from the burning material. The heat shield is also designed to
not burn away completely through the ablation process. There will be some remaining solid
material on the structure to aid in the protecting of the spacecraft from superheated gases. It has
been used on spacecraft since the Apollo era and was going to be used on the new Orion CEV.
The largest benefit to using Avcoat is that it has not been used on any missions that have failed,
so it has a long and unblemished record of accomplishment with the United State’s Space
Program.
Detailed Vehicle Descriptions Crew Capsule Page 359
Author: Alexander Roth
The specific type of Avcoat we select is Avcoat 5026-39H/CG. Table 5.5.7.3-3 below lists its
material properties.
Table 5.5.7.3-3 Avcoat 5026-39H/CG material properties [4]
Property Value Units
Density 5290 kg/m3
Thermal Conductivity (Isotropic) 0.24 W/m-K
Specific Heat 1610 J/kg-K
Emissivity 0.67
Combustion Enthalpy 2.76E7 J/kg
Heat of Vaporization 2.65E7 J/kg
Heat of Decomposition 1.16E6 J/kg
Failure Mode Char spall
This produces a heat shield with a constant layer of Avcoat of 19.2096 kg with an ablative
surface thickness of 0.15 cm. A section view of the heat shield’s structure in Fig. 5.5.7.3-4
shows the layers that compose of the thermal protection system (not to scale).
Figure 5.5.7.3-4 shows the three layers of the thermal protection system of the Capsule
(not to scale). The first outside layer is Avcoat, with a thickness of 0.15 cm, as shown in dark
red. The next layer is insulation, which our code does not compute, but it often is twice the
ablator thickness, so it is 0.30 cm thick, as shown in pink. Finally, the last layer is steel outer
wall, shown in grey. This outer wall protects the contents of the inside of the capsule separate
from the harsh outside and is 66.703 cm thick. In addition, this curved area will host storage for
the astronauts below their feet while inside. Table 5.4.10.2-2 shows the calculated thickness of
Avcoat that the aeroshell will minimally need for the aero maneuvers to work correctly and
safely.
Table 5.5.7.3-4 Thickness and Mass of Avcoat 5026-39H/CG for Capsule
Thickness, cm Mass , kg
0.15 46.765
Detailed Vehicle Descriptions Crew Capsule Page 360
Author: Alexander Roth
Figure 5.5.7.3-4 Section view of heat shield showing the TPS material layers (Avcoat,
Insulation, and Support Structure)
0.15 cm
0.30 cm
~66.703 cm
Detailed Vehicle Descriptions Crew Capsule Page 361
Author: Alexander Roth
References
[1] NASA, “NASA Exploration Systems Architecture Study – Final Report,” NASA-TM-
2005-214062, November 2005.
[2] Davies, C., “Planetary Mission Entry Vehicles Quick Reference Guide, Version 3.0,”
NASA/SP-2006-3401, ELORET Corporation.
[3] Graves Jr., R.A., and Witte, W.G., “Flight-Test Analysis of Apollo Heat-Shield Material
Using the Pacemaker Vehicle System,” NASA TN D-4713, August 1968.
[4] NASA, “TPSX Web Edition V4,” Material Properties Database Web Edition V4.3 [online
database], URL: http://tpsx.arc.nasa.gov [Accessed 02/17/11].
Detailed Vehicle Descriptions Crew Capsule Page 362
Author: Justin Axsom
5.5.8 Communication Systems
The Crew Capsule starts communicating with the Crew Transfer Vehicle as soon as it
jettisons from the Crew Launch Vehicle. We employ the use of an ultra-high frequency, phased-
array antenna to transmit and receive signals to and from the CTV. An image of the antenna
appears in Fig. 5.5.8-1 and appendix A-5.5.8 provides more information on the phased-array
antenna design.
Figure 5.5.8-1 A model of the phased-array antenna. Each of the circular units covers the
individual antennas as described in appendix A.5.5.8.
We use the link to aide in the autonomous docking procedure between the CTV and the crew
capsule in addition to transmitting data. A hatch protects the antenna from the elements during
re-entry and launch, and then extends during operating conditions. During launch, the crew
launch vehicle provides the communication link to mission control, but during re-entry there is a
communication black-out time until it reaches a lower velocity and the plasma around the crew
capsule subsides. At this point, a small omnidirectional antenna serves as a location beacon for
By: Alexander Roth
Detailed Vehicle Descriptions Crew Capsule Page 363
Author: Justin Axsom
crew recovery. Table 5.5.8-1 summarizes the design parameters for the crew capsule
communication system. The link provides a sufficient amount of bandwidth to transmit two
HDTV signals and all logistics data. This data relays through the CTV to the NASA tracking and
data relay satellites which transmit and receive all data to and from mission control.
Table 5.5.8-1 Design Parameters of the Crew Capsule to CTV Link
Property Value
Frequency, GHz 1.20
Data Rate, Mbps 50
Transmitter/Receiver
Power, kW 0.20
Mass, kg 13.5
Diameter, m 0.50
Pointing Range, deg 120
Peak Gain, dBi 28.8
Detailed Vehicle Descriptions Crew Capsule Page 364
Author: Paul Frakes
5.5.9 Capsule Operations
The crew spends a relatively short amount of time in the Crew Capsule. We launch the crew
in the Capsule and then dock the Capsule with the Crew Transfer Vehicle (CTV), at which point
the crew exits the Capsule into the CTV. This process occurs in Low Earth Orbit (LEO). The
Capsule remains unmanned throughout the mission until the crew returns to Earth, at which point
they board the Capsule and enter the atmosphere. The Capsule lands in the ocean and we recover
the crew. The Capsule is also recovered, which is the end of its operational life.
We require the Crew Capsule to perform six separate maneuvers during its operational life.
The first maneuver is the initial docking maneuver with the CTV for crew transfer. The next four
maneuvers entail undocking and subsequent re-docking with the CTV at different locations. We
require these maneuvers (a) to keep the CTV nearly axisymmetric while it spins, thus alleviating
some CTV attitude stabilization concerns, and (b) to allow easy access to the regolith storage
compartment on the Capsule during Ceres surface operations. We assume for these four near-
CTV maneuvers that the CTV is fixed in inertial space, since the maneuver time is small
compared to the period of LEO or Low Ceres Orbit (LCO). Finally, upon Earth arrival, the Crew
Capsule detaches from the CTV and proceeds with atmospheric entry. If any single Crew
Capsule maneuver fails, the entire mission fails and the crew is lost.
Detailed Vehicle Descriptions Crew Capsule Page 365
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5.5.9.1 Rendezvous and Docking to CTV
We insert the Crew Capsule into the same orbit as the CTV, which is already in LEO at an
altitude of 350 km. We launch the Crew Capsule out of phase with the CTV to avoid collision.
The Crew Capsule’s attitude thrusters then perform a burn to slow down and drop the Capsule
into a slightly lower orbit, allowing it to “catch up” to the CTV (i.e., eliminate the phase
difference between the orbits of the two vehicles). As the Crew Capsule nears the CTV, it again
uses its attitude thrusters to raise its orbit to meet the CTV. During this process, the Crew
Capsule performs small, impulsive burns with its attitude thrusters to maintain orientation and to
slowly approach the CTV. The maximum relative speed between the two vehicles is 5 m/s during
the entire docking process. We define the maximum distance for “close proximity operations” of
spacecraft to be 1 km [1]. When the distance between the CTV and the Capsule is less than this
prescribed 1 km, the maximum relative speed between the two vehicles is 0.3 m/s [2]. This
docking procedure is an automatic process, and we have included a backup system to allow for
manual override if necessary. The manual backup system consists of a joystick and computer
control system. Automatic docking processes are well-established and currently in use, for
example, in the docking of the ATV (Automated Transfer Vehicle) with the International Space
Station. Therefore, we do not anticipate any major difficulties executing this procedure.
Once docked, the crew transfers through the hatch in the top of the Crew Capsule into the
crew quarters on the CTV, as shown in Fig. 5.5.9.1-1.
Detailed Vehicle Descriptions Crew Capsule Page 366
Author: Paul Frakes
Figure 5.5.9.1-1 Crew Capsule (green and black) shown docked to the crew quarters on the
CTV (light blue)
By: Alex Roth
Detailed Vehicle Descriptions Crew Capsule Page 367
Author: Paul Frakes
5.5.9.2 Earth Departure
Shortly before the CTV/Crew Capsule assembly departs from LEO, the Crew Capsule will
undock from the crew quarters and perform several close proximity maneuvers to move to the
top of the CTV configuration, labeled (1) in Fig. 5.5.9.2-1. We reposition the Capsule to keep the
CTV configuration nearly axisymmetric as it spins during major burns and interplanetary transfer
phases of the mission, alleviating some CTV attitude stabilization concerns.
Figure 5.5.9.2-1 Proximity operations of the Crew Capsule near the CTV. (1, 3) Capsule
docks to the top of the CTV, and (2, 4) Capsule docks to the crew quarters.
By: Alex Roth and Paul Frakes
Detailed Vehicle Descriptions Crew Capsule Page 368
Author: Paul Frakes
We require 10 separate burns of the Capsule attitude control system during this procedure,
which correspond to five maneuvers. Each of the five maneuvers requires a “speed up” burn and
a “slow down” burn, resulting in the 10 total burns. We show these five maneuvers in Fig.
5.5.9.2-2.
Figure 5.5.9.2-2 Simplified schematic of the five individual maneuvers (10 burns) required
to move the Crew Capsule (blue) from one docking port on the CTV to another.
The first maneuver (labeled (a) in Fig. 5.5.9.2-2) separates the Crew Capsule from the CTV.
The second (b) spins the Crew Capsule 90 degrees, orienting it in the proper direction for re-
docking. The third maneuver (c) translates the Crew Capsule along the length of the CTV, past
the aft end (the end opposite the crew quarters). The fourth maneuver (d) axially aligns the Crew
Capsule with the CTV. The fifth and final maneuver (e) translates the Capsule and docks it with
the CTV, as shown in Fig. 5.5.9.2-2. All of these maneuvers are automatic. In the event that the
By: Paul Frakes
Detailed Vehicle Descriptions Crew Capsule Page 369
Author: Paul Frakes
automatic control system fails, we require remote manual control of the system because the Crew
Capsule is unmanned during this maneuver. A joystick and computer serve as the remote control
system, and we place this system in the crew quarters of the CTV.
After the Crew Capsule is secured once again to the CTV in its new location, the CTV will be
ready for LEO departure.
Detailed Vehicle Descriptions Crew Capsule Page 370
Author: Paul Frakes
5.5.9.3 Ceres Arrival
Upon arrival at Ceres, we employ the steps described in Section 5.5.9.2, in reverse order, to
return the Crew Capsule to its position on the crew quarters of the CTV. This maneuver is
labeled (2) in Fig. 5.5.9.2-1. We perform these maneuvers in LCO, prior to Ceres descent, but
after the CTV de-spins. The repositioning of the Crew Capsule will allow the crew to have easy
access to the regolith storage compartment when the CTV is on the surface of Ceres.
Detailed Vehicle Descriptions Crew Capsule Page 371
Author: Paul Frakes
5.5.9.4 Ceres Departure
We must reposition the Crew Capsule before Ceres departure to again ensure the CTV
configuration is nearly axisymmetric for the return transfer. Similar to the previous maneuvers,
we again employ the steps described in Section 5.5.9.2 to return the Crew Capsule to its position
at the aft end of the CTV. The reposition maneuvers occur in LCO and are again autonomous.
This maneuver is labeled (3) in Fig. 5.5.9.2-1.
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5.5.9.5 Earth Arrival and Atmospheric Re-Entry
When the CTV de-spins and arrives at Earth, the Crew Capsule again follows the procedure
described in Section 5.5.9.2, in reverse order, to return the Crew Capsule to the crew quarters on
the CTV. This maneuver is labeled (4) in Fig. 5.5.9.2-1. At this time, the crew enters the Capsule
and prepares for departure and atmospheric entry. The Capsule then undocks and performs one
last small burn to separate it from the CTV, so that a safe distance is maintained between the two
vehicles as each ballute is deployed. The Crew Capsule deploys its ballute to slow down through
the upper atmosphere and then employs parachutes to land safely in the ocean.
Detailed Vehicle Descriptions Crew Capsule Page 373
Author: Paul Frakes
References
[1] Carrico, T., Langster, T., Carrico, J., Vallado, D., Loucks, M., and Alfano, S., “Proximity
Operations for Space Situational Awareness,” Advanced Maui Optical and Space
Surveillance Technologies Conference, 2006.
[2] Noll, R. B., Zvara, J., and Deyst, J. J., “Spacecraft Attitude Control During Thrusting
Maneuvers,” NASA Marshall Space Flight Center, 1971.
Detailed Vehicle Descriptions Crew Capsule Page 374
Author: Andrew Curtiss
5.5.10. Storage and Return of Ceres Rock
One of the initial project requirements is to return one metric ton of Ceres regolith to Earth. In
order to prevent the contamination of astronauts with the Ceres rock, the regolith container is
isolated. In addition, the container features two separate compartments, one assigned to each of
the ISPP station locations. The first location produces 500 kg of rocks and the second location
produces the same. These samples are separated from each other in case the makeup of the rock
is different on different sides of Ceres. The specific design specifications of the rock return
container are explained in section 5.5.6.2.
To keep the regolith cryogenically frozen for the trip back to Earth, we keep the regolith
container with the stored cryogenic propellant. When the Crew Capsule arrives back in LEO, we
move the compartment into the Crew Capsule for reentry. The insulation of the container keeps
the regolith at cryogenic temperatures for the relatively short flight from LEO to the surface of
the Earth. These measures eliminate the necessity of having a large complex cryogenic
refrigeration system on the Crew Capsule.
Detailed Vehicle Descriptions Crew Capsule Page 375
Author: Devon Parkos
5.5.11. Aerodynamic Maneuvers
v∞ = Vehicle velocity relative to Earth
Δv = Velocity change
σ = Standard deviation of atmospheric density
5.5.11.1. Aerocapture Maneuver
The Capsule aerocapture maneuver is similar to the CTV’s trajectory, but the initial pass is
through a higher altitude of 92.10 km. Before arriving at earth, the Capsule departs from the
CTV, performing a small burn that shifts its perigee altitude. V∞ remains at 7.89 km/s. The
relatively more significant effect of the Capsule’s aerodynamics on the trajectory causes this
behavior. We still constrained the elliptic orbit to remain outside the moon’s sphere of influence
and not expose the vehicle to more than 9 g’s.
Our Capsule trajectory for 2.5 standard deviations of atmospheric density in either direction is
shown below, in Fig. 5.5.11.1-1. By considering all possibilities in this range, we maintain the
same 98.8% success rate. While the apogee altitude varies between the CTV and Capsule
maneuvers, the elliptic behavior afterwards is extremely similar. We desire the same Δv for each
vehicle (the minimum value to meet our successful capture criterion), which causes the trajectory
limits to align. In actuallity, the experienced atmospheric densities for each trajectory could be
very different, and the Capsule and CTV can end up on opposite ends of the uncertain trajectory
range.
Detailed Vehicle Descriptions Crew Capsule Page 376
Author: Devon Parkos
Figure 5.5.11.1-1 The deviation in trajectories for the Capsule after capture.
The largest possible elliptic orbit considered corresponds to a maneuver time of 7.6 days. We
must provide supplies for at least this amount of time. The minimum maneuver time is
approximately 6 hours. We sized the heat shielding on the vehicle to withstand both the longest
ablation times and the peak heating rates, which occur during opposite ends of the atmospheric
density uncertainty. The ablative layer thicknesses for each case are examined in section 5.5.7.
Detailed Vehicle Descriptions Crew Capsule Page 377
Author: Devon Parkos
5.5.11.2. Atmospheric Re-entry and Landing
After the aerocapture maneuver, the Capsule’s trajectory deviates further from the CTV’s
path. The Capsule retains its ballute for a longer time and greater Δv, which results in the
Capsule remaining within the denser range of the atmosphere. By timing the release, similar to
the CTV, we mitigate the effect of the atmospheric uncertainty prevalent in the upper
atmosphere. By choosing the release time, our trajectory will result in a landing location that is
less dependent on the atmospheric density experienced. This allows us to choose a return time
that will result in a safe landing location. The trajectories for 2.5σ uncertainty values in either
direction are shown below.
Figure 5.5.11.2-1 The capsule trajectories near Earth.
Detailed Vehicle Descriptions Crew Capsule Page 378
Author: Devon Parkos
Figure 5.5.11.2-2 The landing behavior and location relative to initial conditions.
Shortly after the ballute releases, we deploy the Capsule parachutes and slow to terminal
velocity. Slight ablation occurs, as visible in the previous figures (See section 5.5.7). We sized
the parachutes to remain safe even with a single chute failure. As can be seen in the above
figure, our Capsule descent is slowed to a terminal velocity, verifying the behavior we expected
from sufficient parachute sizing. More information on the descent is shown in figs. 5.5.11.2-3
and 5.5.11.2-4.
Detailed Vehicle Descriptions Crew Capsule Page 379
Author: Devon Parkos
Figure 5.5.11.2-3 Altitude profile for Capsule in high density case.
Figure 5.5.11.2-4 Velocity profile for Capsule in high density case
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5.6. In-Situ Propellant Production Stations
The following section discusses the In Situ Propellant Production (ISPP) stations. These
stations produce all of the hydrogen, oxygen and water required to safely complete our mission.
5.6.1. ISPP Configuration
Figure 5.6.1-1 The ISPP station in full extension after landing and the regolith collection
operation has begun
We initially place the ISPP stations inside the Supply Transfer Vehicle’s (STV) central
module. We store the stations’ Harvesters inside the cargo modules onboard the STV. Figure
5.6.1-1 reveals the ISPP station in full extension or operation mode. The central module for the
ISPP station consists of a nuclear reactor, a water extraction oven, an electrolizer, condensers,
pumps, power conversion system, radiators, and a multitude of pipes. We reuse the liquid
hydrogen and liquid oxygen tanks from the STV. These tanks (not pictured) attach to the central
module seen in Fig. 5.6.1-1 along with a separate water holding tank (light blue tank). The
By: Alex Roth
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radiators unfold from their condensed box form into the five square set observed in the figure
above. Upon arrival, we deploy the conveyor belts and the regolith collection bin as seen in the
following figure:
Figure 5.6.1-2 The ISPP station exhibiting the collection bin (left), input conveyor belt
(middle), oven (red), and the output conveyor belt (below oven on left)
The Harvesters deploy from their corresponding STV modules as the modules rest on the
surface after landing. From here they begin the collection of regolith around the ISPP station,
collecting the regolith into the collection bin at a rate of 62 tons per day. The configuration of the
Harvesters is found in Section 5.6.4.1. The ISPP operates for 2.256 years and produces the
amount of water, hydrogen, and oxygen displayed in the table below:
By: Alex Roth
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Table 5.6.1-1 ISPP Production Values
Total Production Values
(for 1 ISPP station)
Values Units
Production Time 2.256 Years
Production Time 824 Days
*Water extracted 46.39 T
**Hydrogen extracted 118.5 T
**Oxygen extracted 545.6 T
*Stored at Ceres ambient temp (ice)
**Stored in liquid form
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5.6.2. ISPP Production Timeline
In the year 2025, the In Situ Propellant Production stations depart for Ceres onboard STV-1
and STV-2. The ISPP station operations begin shortly after landing on Ceres. After receiving the
start up signal from Earth, the reactor takes approximately 11 hours to heat up and liquefy the
reactor coolant. Once the power system activates, we have the ISPP stations and Harvesters
begin production operations. The Harvesters make four regolith collection trips each day that
take three hours per trip. The Harvesters use the remaining downtime for battery regeneration
and for clearing away the excess “used” regolith. The process takes approximately 2.256 years at
a rate of 63 tons of regolith collection per day for all production to be completed. Upon
completion of propellant production the reactor only supplies energy to the liquid hydrogen and
oxygen tanks to keep them at the necessary temperatures to maintain a liquid state. We then use
the reactor to reheat the water into a liquid state following the crew’s arrival. The station then
resupplies the Crew Transfer Vehicle with necessary propellant, water and breathable oxygen
within 24 hours of its arrival. The Rovers treat the ISPP stations as a refueling depot where they
go to replenish any depleted propellant, oxygen, and water stores used on missions. The crew
uses the first station for 196 Earth days before moving to the second station for the next 196
Earth days. Upon the crew’s departure from a station, that station powers down operations. We
placed the graphical representation of the ISPP production life in the following image.
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Authors: Matthew Hill, Zachary Richardson
Figure 5.6.2-1 The ISPP flowchart for one station from Ceres landing to end of life
The green boxes denote STV or CTV actions while the blue represent key points in the ISPP
production. The end of life configuration is discussed in more detail in Section 5.6.8.
STV-1 & 2 Launch and Assembly in
Orbit
STVs Transfer to Ceres
STVs Entry, Descent and
Landing
Station setup, Begin Harvester
OPS
Propellant Production Cycle
(~2.3 Years)
Secure Harvesters as Reactor powers propellant holding
tanks
CTV Transfers to Ceres (Station
maintains Tanks)
CTV Arrival, propellant
transfer
Support surface operations (Rover
refueling)
Crew transfers to Station 2 (ISPP-1 powers down)
Station 2 repeats refueling ops for
CTV/Rovers
Crew Departs after 392 days total on Ceres
End of Life Ops
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5.6.3. ISPP Power Systems
5.6.3.1. ISPP Reactor
Over the course of this project we considered and evaluated numerous reactor design concepts
for the ISPP stations and were forced to revise our selected design several times as our
understanding of the requirements for the propellant extraction process and for surface
operations evolved. Our initial design selections were based on pressurized water reactors; in
retrospect these designs are far too massive and would have also produced much more power
than our project requires. Other concepts evaluated included emerging designs for small,
“modular” reactors like the Hyperion Power Module, which produces thermal power in the range
of 20MWth from a reactor vessel of impressively small size (circa 20 T and 3 m3). These
systems are also too powerful for use in our spacecraft. Ultimately we have been more successful
when adapting designs intended for use in zero-gravity than for operations on earth.
The design which we ultimately selected for development is the heat-pipe reactor, so-called
because its core fuel rods are clustered in threes around heat-pipes containing the molybdenum-
sodium alloy coolant, which transports heat generated by the reaction from the core to a gas heat
exchanger in the stirling engine (see Section 5.6.4.2). Scaling from the existing SAFE 400 design
(by Los Alamos National Laboratory) the reactor for this mission features a core radius of 24cm,
a core length of 133cm and will has a core mass of 5.243T [1]. The reactor produces 2.2 thermal
Megawatts of power, which will be split between 260 thermal Kilowatts to the extraction oven
(supplied directly through a heat exchanger) and 680 electrical Kilowatts for the remaining
systems (supplied through the power conversion system).
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Figure 5.6.3.1-1 this image shows the reactor vessel in relation to the other ISPP station
machinery (radiation shielding around the reactor is not shown); the reactor connects via
heat pipes to the oven (bright red, right) and to the radiators (not shown)
Like its progenitor, the SAFE-400, the ISPP reactor uses uranium nitride (UN) as its primary
fuel; this is assembled in an array of 456 rhenium-clad fuel pins [1]. The reactor vessel, the
structure that contains the core and its heat collection systems, is constructed from Hafnium,
which is used because of its excellent ability to capture neutron radiation (600 times greater than
zirconium; see section 5.6.4.4 for discussion of shielding) and because it withstands extreme
temperature conditions well. The reactor operates at a temperature of 1020 degrees Celsius under
normal conditions but is capable of exceeding this under emergency conditions.
By:Alex Roth
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Author: Matthew Hill
5.6.3.2. ISPP Power Transfer Systems
While our water draws thermal power directly from the reactor heat exchange system, the
majority of the systems in the ISPP stations require electrical power, meaning that the design
must include a power conversion system. In the ISPP stations we have selected a stirling engine
for this purpose. Stirling engines are capable of very high levels of efficiency, approaching 40%;
for the purpose of this design exercise we assumed an efficiency of 35%, which we anticipate
would be realistic for a large stirling engine [4]. When coupled with the turbo-alternator this
system produces electricity with an efficiency of approximately 30%. The combined mass of the
Stirling engine-alternator system is 5.55 T, with a volume of 19.6 cubic meters. The stirling
engine is connected to the alternator via a driveshaft.
Figure 5.6.4.2-1 this cutaway image of the ISPP machinery shows the position of the
Stirling Engine module (deep red, bottom center) and the turbo-alternator (tan, bottom
right) in relation to the other components. In its actual configuration this machinery would
be obscured by the hull of the STV center module
Thermal energy is conducted from the reactor to the oven via a separate heat pipe system that
connects to a set of heating coils embedded in the walls of the oven (this heat pipe is visible in
By: Alex Roth
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 389
Author: Matthew Hill
Fig. 5.6.4.2-1). The thermal power provided by the reactor far exceeds the thermal power
required for the oven (260 kW); The reactor is theoretically able to operate all systems at
maximum power simultaneously. We do not anticipate that this will be necessary but we have
designed the system to be capable of providing this much power as a contingency. If necessary
the reactor can also operate at temperatures higher than its rated operating temperature for short
periods of time [3], however this is not advisable as the additional heat buildup may damage
other components within the spacecraft.
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5.6.3.3. ISPP Heat Dissipation Systems
An important consideration for a spacecraft design that produces enormous amounts of waste
heat such as this one is how that heat will be rejected from the spacecraft. Since we were unable
to verify the utility of the planet’s surface as a heat sink we were left with a more conventional
method of heat dissipation, the radiator. While we considered several more exotic radiator
designs (such as using a flowing sheet of liquid as a heat-exchange surface), the ultimately
selected design is based on the one developed for use with the same SAFE-400 reactor that our
own reactor is based on [6]. This radiator design consists of composite panels made from carbon-
carbon plumbed with heat pipes (which carry the radiator coolant), which in the original design
is a mixture of helium and xenon. In its original incarnation the radiator design was capable of
rejecting 400kWth. For the purposes of the ISPP station the radiator is required to reject
approximately 2.3 MWth assuming that no work was being performed and the reactor was
operating at maximum power (which most likely would not take place under normal conditions).
This necessitated some modifications to the original design. Using the equation:
(5.6.3.3-1)
Where is the Stefan-Boltzman constant, is the emissivity of Carbon-Carbon, T1 is the input
temperature of the radiator coolant, T2 is the ambient temperature on Ceres, and Q is the rate at
which energy must be rejected (this should account for solar radiation flux), then the area A was
found to be 83.3 square meters, which we rounded to 95 square meters to provide a margin of
safety in the event that the reactor exceeds its rated operating temperature (this is possible if
coolant flow is interrupted). The specific mass of the SAFE design is 14 kg/square meter (which
will not change since the composition of the panels is unchanged) [1], meaning that the mass of
the radiator area required would be 1.33T. The design of the radiators features five panels (each
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with an area of 19 arranged in a cruciform pattern, with the four outer panels folding onto the
center panel during transit from earth. On Surface Operations Day 1, these panels unfold and the
reactor starts, with heat being slowly increased in order to melt the coolant in the radiator system
(which will harden in low temperatures). It is estimated that this process will require 11 hours
[1].
Figure 5.6.3.3-1 Radiators in their deployed configuration; for transportation to Ceres the
four outer panels will fold onto the center panel to create a more compact shape
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5.6.3.4. ISPP Radiation Protection and Mitigation
The principal forms of ionizing radiation with the potential to be hazardous for the crew
during surface operations on Ceres are neutron radiation (consisting of liberated neutrons
produced in the course of the fission reaction) and gamma (γ) radiation (high-energy
electromagnetic radiation).
Exposure to radiation can be mitigated through several measures, including adding radiation
shielding to either the source of emission or the recipient, by increasing the distance between the
emission source and the recipient, and by limiting the time that the recipients spend in contact
with the radiation source. The constraints of the mission profile, which require that the crew
operate on the surface for a year and that their habitat land close enough to the ISPP station to
facilitate transfer of consumables between the two spacecraft limits the usefulness of two of
these strategies. As a result, our primary means of radiation protection necessarily became the
addition of shielding to the ISPP station.
Several studies conducted by various institutions, including NASA, the Nuclear Regulatory
Commission and the Department of Energy have investigated the utility of using soil and rock
recovered from the surface to construct a barrier against radiation. These studies suggest that this
would be a viable option depending on the mass- and linear attenuation coefficient of the
materials used. Generally, this could be accomplished either by creating a trench or pit in the
surface of Ceres and lowering the entire reactor plant into this hole, or by excavating materials
from the surface and using them to construct a wall around the reactor. While the geometry of
the ISPP stations does not permit the first option, the nature of ISPP operations, which require
our harvesters to excavate and then shift large amounts of regolith make our system ideally
suited to the application of the second method. Using the regolith cast off from the water
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extraction process the harvesters could construct a semi-circular wall on the side of the supply
transfer vehicle facing the CTV landing site, protecting the CTV once it arrives. However, this
approach faces several hurdles. The first of these is the fact that little is known about the physical
composition (especially attenuative properties) of Ceres or other objects in the Asteroid belt.
Since the dimensions of the barrier that must be constructed are primarily based on how well the
Ceres regolith absorbs gamma and neutron radiation, this makes assessing the practicality of this
approach difficult. As a result we used the mass attenuation coefficient and linear attenuation
coefficient of Lunar regolith as an analogue for Ceres regolith.
As future reconnaissance missions to Ceres return information about the composition of the
surface there we will be able to revise the dimensions of our planned radiation barrier. Since the
crew would need to pass inside any radiation barrier for long enough to collect supplies from the
STV cargo containers, their exposure to radiation would most likely be very high if they were
required to perform this task in the presence of an unshielded reactor. This makes the application
of shielding directly to the reactor machinery necessary in addition to the shielding constructed
on the surface.
Since our primary concern is attenuating the two primary forms of ionizing radiation
produced by this reactor we have selected a relatively simple four-layer shield composed of three
materials: Lithium Hydride (LiH), Boron Carbide (B4C) and Tungsten (W). LiH was selected
for its ability to attenuate neutron radiation and for its very low density (.82 g/cubic cm)[2]. The
innermost layer of LiH is approximately 10cm thick and will serve as a low-z buffer to the
second layer of shielding, 8cm of Tungsten, which is a powerful gamma attenuator. Due to the
high density of tungsten only a thin layer is used. The third layer of shielding is also composed
of LiH, this time a thicker layer 80 cm in width. The outermost layer of the shielding is made
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from Boron Carbide, which has the advantage of being an effective neutron and gamma
attenuator as well as being an extremely hard ceramic that is often used for ballistic protection in
body armor and military vehicles. As a result this shielding layer will provide effective
protection to the reactor from environmental hazards such as micrometeoroids and condensation.
The layer will be 2cm thick.
Figure 6.6.4.4-1 Cross-sectional image showing the layers of radiation shielding on the
reactor; the inner red circle represents the core (hafnium); the first circle is LiH, the next is
Tungsten, the next is also LiH. The outer black circle is B4C. Figure is not shown to scale
for the purpose of making the thinner layers visible
The total shielding thickness is 95cm with a mass of 2269.8 kg. With this shielding alone a crew
member receives a radiation dose of approximately 50 rem/year at a range of 150m from the
reactor (this is the standard maximum dosage rate set by NASA)[2]. Crew operations are
structured to minimize the amount of time they are in close proximity to the reactor; plans call
for the crew to approach within a 20 meter exclusion zone for only short periods of time (less
than 1.5 hours at a time). In order to bring the minimum range of safe dosage closer to the
landed supply transfer vehicle the ISPP harvesters will be capable of constructing an additional
M. Hill
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radiation barrier from discarded regolith (cast off from the water extraction process). Based on
the mass attenuation constant of lunar regolith a semicircular barrier 3m in height and 2.5m thick
at a distance of 20m from the central module of the supply transfer vehicle (containing the
reactor) would result in a dosage of 50 rem/year at a range of 30m within a radioactive “umbra”
obscured by the wall [2]. This analysis does not account for the attenuative properties of the
spacesuits and the hull of the Crew Transfer Vehicle, which will not be insignificant and will
further protect the crew from radiation produced by this reactor.
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5.6.3.5. ISPP Power Operations Profile
Following landing on Ceres the ISPP station awaits a command to begin operations. Upon
receipt of this command the station deploys its radiator and auxiliary power is activated,
providing power to start the reactor. Since during transit the reactor is not running the coolant in
the heat exchanger for the radiator will have solidified in the low temperatures. The reactor
operating temperature must be increased slowly in order to liquefy this coolant. The startup and
burn-in process requires approximately 11 hours and ends when the reactor reaches peak coolant
temperature (1800 K). At this point the station begins operations. As water is extracted and the
electrolysis and storage systems come online the power demand on the reactor increases; at
maximum operational tempo the reactor operates at 90% capacity. Once all the requisite water
has been collected (after approximately 2.3 years) the station shuts down with the exception of
providing power to cooling units in the propellant storage tanks to maintain the propellant (LOX
and LH2) in liquid form.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 397
Author: Matthew Hill
References
[1] Bushman, A., Carpenter, T., & Ellis, S. e. (2003). The Mars Surface Reactor (MIT-NSA-
TR-003). MIT.
[2] Craft, A. E., & King, J. C. (2009). Radiation Shielding Options for the Affordable Fission
Surface Power System. Rolla: Missouri University of Science and Technology.
[3] Mason, L. S. (2003). A Power Conversion Concept for the Jupiter Icy Moons. Hanover,
MD: NCIA.
[4] Nightingale, N. P. (1986). Automotive Sterling Engine Mod II Design Report. Cleveland,
OH: NASA Lewis Research Center.
[5] Oi, T., & Sakaki, Y. (2003). Optimum hydrogen generation capacity and current density of
the PEM-type water electrolyzer operated only during the off-peak period of electricity
demand. Amsterdam: Elsevier.
[6] Vaughn, W., Shinn, E., Rawal, S., & Wright, J. (1998). Carbon-Carbon Composite
Radiator Development for the EO-1 Spacecraft. Hampton: NASA Langley Spaceflight
Center.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 398
Author: Elle Stephan
5.6.4. Harvesters Detailed Description
5.6.4.1. Harvester Configuration
Each Harvester incorporates a box-like cargo frame made of Aluminum 7075. The cargo
frame encases the Li-ion battery and electric motor as well as the small computer chip necessary
to operate the communications antenna. The battery, motor, and communications system are
housed in the back section of the main cargo frame. A small camera is mounted on the front,
upper surface of the Harvester to maintain constant visual communication before the astronauts
reach Ceres. Below, Fig. 5.6.4.1-1 gives the basic dimensions of the Harvester including the
drive system.
Figure 5.6.4.1 - 1 The Harvester implements a rocker-bogie drive system, seen with black
wheels and orange legs
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 399
Author: Elle Stephan
The Harvester’s main structure holds one fourth of the total regolith collected in a 24hr day
with a volume of 7.53m3. The compact volume of the regolith is just over 3.5m
3; however, we
assume the regolith does not pack perfectly into 3.5m3 cubes. The volume difference allows for
40% of “empty space” between the rocks of regolith.
The drive system incorporates a rocker-bogie type structure (pictured in Fig. 5.6.4.1-1) that
allows for ease of movement over uneven terrain. Each wheel on the rocker-bogie system
operates independently of the other wheels by mounting the wheel supports on two separate
joints seen in the above figure. The wheels have a diameter of 0.25 meters and a thickness of
0.25 meters for suitable surface traction.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 400
Author: Elle Stephan
5.6.4.2. Harvester Power Systems
One Li-ion battery powers each Harvester. The battery recharges for three hours on a weekly
basis by a connection to the nuclear reactor powering the ISPP station. The battery generates
power for the communication system, the motor, and largely the camera.
Table 5.6.4.2 - 1 Power requirement component breakdown of for each Harvester
Power Required
Communication System 0.1 W
Motor 19.2 W
Camera 1560 W
Total Power 1579.3 W
The communication system and motor power requirement is substantially lower than the
camera power requirement because the communication transmitted is basic data (explained
further in Section 5.6.4.7) and the Harvester does not reach large enough speeds to warrant the
motor to produce large amounts of energy.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 401
Author: Elle Stephan
5.6.4.3. Harvester Propulsion Systems
A combination of a battery and an electric motor attached to a six-wheeled rocker-bogie
system propels the Harvester on it’s over two yearlong mission. As explained in Section 5.6.4.2,
a Li-ion rechargeable battery provides power to the motor that transfers electric power to the
19.2W of mechanical power needed for the rocker-bogie.
Three wheels of equal size (0.25m in diameter by 0.25m thick) equip each of two rocker-
bogies. The front wheels are extended about one meter from the front edge of the Harvester,
which allows for better balance and mobility. The back wheels stand slightly behind the
Harvester’s back end and serve the same purpose as the front. The center wheels support most of
the Harvester weight and provide better balance for the structure.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 402
Author: Elle Stephan
5.6.4.4. Harvester Structural and Thermal Systems
The cargo bay of the Harvester employs Aluminum 7075 with a density of 2810kg/m3 as the
structural material encasing the regolith. The five main panels (top, bottom, back, two sides) and
the half panel on the upper front section all have a thickness of 0.5cm. The bottom front panel
that opens for regolith collection is twice as thick at 1.0cm to create more weight, and therefore,
more friction for collection (explained further in Section 5.6.4.5). Below, a table shows
dimensions of each panel.
Table 5.6.4.5-1 Dimensions of each panel with thickness of 0.5cm excluding the drop down
panel
Sides Front (top panel) Top Bottom Back
Height, m 1.75 0.75 1.75 1.75 1.5
Length, m 2.75 1.75 2.75 2 1.75
Area, m2 2.63 1.31 4.81 3.5 2.65
Table 5.6.4.5-2 Dimensions of drop down panel with 1.0cm thickness
Drop Down Panel
Bottom length, m 1.5
Bottom width, m 1.75
Side height x2, m 1.06
Side width x2, m 0.5
Area, m2 1.94
The sides of the drop down panel fit into the cargo bay when the door is closed. When open,
the sides provide a lip to catch regolith from simply falling off the edges.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 403
Author: Elle Stephan
5.6.4.5. Harvester Collection Process
We made a few assumptions in order to design the collection process. First, we assume that
the Harvesters collect a meter deep of regolith and travel 87.7m in radius from the ISPP station.
These dimensions fulfill the mission’s volume requirement of 24208.5m3. To complete the
Harvester’s purpose in 2.3 years, they must collect a combined volume of 29.3m3 in a 24hr day.
The heavy drop down panel, or door, remains open and digs into the Ceres regolith, shoving it
into the cargo bay until a sensor registers that the Harvesters collect 7.75T of regolith. The door
then closes for the return to the ISPP station where it drops off the regolith into a collection bin.
With a volume capacity of over 3.5m3 and two Harvesters working at a time, each Harvester
makes four trips per 24hr day at a speed of 0.016 m/s. The Harvesters are allowed 12hrs (3hrs
per trip) to collect the 29.3m3 of regolith needed per Earth day allowing another 12hrs of regolith
dispersion with the exception of the 3hr battery recharge period every week.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 404
Author: Elle Stephan
5.6.4.6. Harvester Used Regolith Dispersion Process
The dispersion of the regolith is much similar to the collection only in reverse. The
Harvesters drive through the accumulated pile of baked regolith, filling their cargo bay. They
then, drive previous traveled paths to drop off the baked regolith in a location no longer suitable
for collection.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 405
Author: Sarah Jo DeFini
5.6.4.7. Harvester Communication Systems
Each Harvester is equipped with a low-gain, omnidirectional antenna for communication with
the main ISPP facility. The range of these antennas is approximately 150 meters, which is well
in excess of the required range of 100 meters. See the appendix D.2.1.5 for a detailed
description of the sizing process. Table 5.6.4.7-1 is a summary of the physical parameters for
each antenna.
Table 5.6.4.7-1: Harvester antenna specifications
Length, m 0.0485
Radius, m 0.005
Mass, kg 0.034
Efficiency 0.9997
Gain, dB 1.631
Beamwidth, degrees 124
Radiated Power, kW 0.0409
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 406
Author: Zachary Richardson
5.6.5. Extractor Detailed Description
The “Extractor” refers to all of the components of the In-Situ Propulsion Production station
that play a role in the physical extraction of water and production of propellant. The Extractor
includes the oven/conveyor belt system, the electrolysis machine (electrolizer), the various pipes
connecting each component, the heat pumps and the condensers and the communication and
operating computer. The table below shows each of the Extractor components’ mass, power
required and volume.
Table 5.6.5.-1 ISPP extractor components
The oven does not require power (*) as the thermal energy generated by the reactor heats the
oven to significant temperatures. The stared value is simply how much power would be required
if we did not chose this method. We based the oven upon industrial batch ovens made by
Precision Quincy [1]. The Electrolysis machine we based upon the Hydrogen Generation
machines created by Hogen Hydrogen [2]. Interpolating both designs allows us to use these
devices in the ISPP station. We assume both technologies are “space-ready” at the start of our
mission timeline.
Specifications of Single ISPP Facility
Component Mass ,T Power, kW Volume, m3
Oven 3.82 369* 43.9
Collection Bin & Conveyor Belt System 0.121 0.401 1.01
Electrolysis Machine 6.26 341 15.3
Pipes/Condensers/Pumps 0.638 70.9 1.91
Computer and Communications 0.016 5.401 0.06
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 407
Author: Zachary Richardson
References
[1] NVISION, “EC Walk-in Ovens”, Precision Quincy. 2011. Date accessed: February 17,
2011. URL:http://www.precisionquincy.com/ovens/info/Industrial_Walk-
In_and_Cart/EC_Walk-in_Oven.
[2] Hogen Hydrogen, “Hydrogen Generation Systems”, Proton Energy Systems. 2011. Date
accessed February 26, 2011. URL: www.protonenergy.com.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 408
Author: Zachary Richardson
5.6.5.1 Extractor Configuration
The extractor consists of the regolith-baking oven, the hydrogen-oxygen electrolizer, the
communication and control system, and the connecting pipes, pumps, and condensers. We house
the extractor in the central STV module. This module consists of a 4.4 m radius base with the
afore mentioned extractor components placed on top. A figure of this section of the ISPP facility
shown below also includes the nuclear reactor, sterling engine, and power conversion system:
Figure: 5.6.5.1-1 The extractor configuration including the oven, reactor, electrolizer,
condensers, pumps, communication dish and pipes.
By: Alex Roth
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 409
Author: Zachary Richardson
Figure: 5.6.5.1-2 The extractor configuration focusing on the oven and conveyor belt
input/output system.
The regolith placed into the collection bin by the Harvesters moves into the oven by the input
conveyor belt. Once we extract the water, the regolith drops onto a second, outward conveyor
belt which dumps the baked regolith nearby for the Harvesters to remove. The oven sits on a
stand that allows it to be a meter above the module base. The input conveyor belt must reach
from the surface of Ceres to the top of the oven, approximately 19 m long. The input conveyor
belt uses container segments to keep the regolith from falling off as it travels up to the oven.
Insulated pipes connect the oven to the electrolyzer and the electrolyzer to the pumps and the
condensers. The pipes move the extracted water and propellant into their corresponding tanks
(See Fig. 5.6.5.1-1).
By: Alex Roth
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 410
Author: Zachary Richardson
5.6.5.2 Extraction Process
The extraction process begins with the Harvester’s collecting 62 tons of regolith every earth
day. The Harvesters collect this amount of regolith in four trips each day. Every half day the
input conveyor belt moves the rock collected by the Harvesters in the collection bin into the oven
for heating. Splitting the regolith into two batches halves the amount of energy required to heat
the regolith to the desired temperature of 200 degrees Celsius. The oven extracts the water once
it becomes water vapor. The vapor travels through pipes with lower insulation levels, thus
exposing them to the colder Ceres ambient temperature where the vapor condenses into a liquid.
The liquid water is then either pumped into a water storage tank where it is kept as ice or it
continues into the electrolizer. Once in the electrolizer, the electrolysis process begins as the
addition of energy to the water begins a chemical reaction that generates both hydrogen and
oxygen [1]. See Eq. 5.6.5.2-1:
(5.6.5.2-1)
The electrolysis process that we use has the following energy requirements and specifications:
Table 5.6.5.2-1 Electrolysis specifications for ISPP facilities
Parameter Value Units
Amount of Water into
hydrolysis machine per day
1674 kg
Amount of Hydrogen
produced in one day
148.8 kg
Amount of Oxygen produced
in a single day
1190 kg
Energy required for 1 kg of
water
17630 kJ
Power required 341.6 kW
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 411
Author: Zachary Richardson
After the generation of hydrogen and oxygen, the two elements split into two separate pipe
systems drawn by two separate pumps. The following figure shows our electrolyzer and the
corresponding oxygen and hydrogen pumps and compressors:
Figure 5.6.5.2-1 Water flows into the first pump (top right) and on into the electrolysis
machine where it is converted into hydrogen and oxygen
Upon separation, the oxygen and hydrogen are condensed into liquid form and pumped into
the STV’s holding tanks. The liquid hydrogen and liquid oxygen remain in the tanks, insulated
from the ambient temperature of Ceres and kept at a low enough temperature to prevent a state
change. We account for boil off with a 20% amount of extra propellant produced than required
for the mission. We trust that future technology also gives us the ability to limit boil off losses.
The ISPP station has the ability to keep producing propellant to account for losses as well.
By: Alex Roth
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 412
Author: Zachary Richardson
References
[1] Nave, R. "Electrolysis of Water." Hyper-physics: Thermodynamics 08 JUN-2005. Physics
at Georgia State University. Web. 16 Feb 2011. <http://hyperphysics.phy-
astr.gsu.edu/hbase/thermo/electrol.html>.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 413
Authors: Zachary Richardson
Co-Author: Leonard Jackson
5.6.5.3. Extractor Structural and Thermal Systems
The extractor structural subsystem in the ISPP station [APES] consists of the oven and the
electrolysis conversion system. The oven’s purpose is to take in the regolith from Ceres and
extract the water from the rocks by heating them. The water vapor is then injected into the
electrolysis conversion system where it is decomposed into hydrogen and oxygen substrates.
The oven attaches on the central module’s base by a set of quad landing legs that account for
any shock absorption on impact. The legs consist of four parts: upper leg, lower leg, and lander
paw. Using code provided by Andrew Curtiss, we were able to design the oleos to lift the oven a
meter off of the modules surface, and come up with masses and volumes.
Table 5.6.5.3-1 Lander Leg Masses and Volumes for All 4 legs
Mass (kg) Volume (m3)
Lander Legs 35.35 0.019
Aluminum alloy (with 4.4% Copper) makes up the majority of the oven. This material is
chosen due to its relatively high heat capacity, low density, and high melting temperature (stats
can be found in the appendix). We decided to overdesign the tank thickness to be 1cm rather than
the 0.001cm thickness a hoop stress analysis predicted. The overdesign is due to the fact that the
oven will be holding several tons of regolith at a time, and by increasing the thickness of the
tank, there will be less of a chance of rupture when the regolith is dumped into the tank.
All of the extractor components such as the oven and pipes are insulated with 50cm of rigid
closed cell polyurethane [1]. We assume the oven and electrolyzer are both space worthy and
their complicated internal components work in micro-gravity. The pipes and conveyor belts
consist of Carbon-Fiber which drastically decreases the mass and increases the strength. The
conveyor belts are based on Carbon-Fiber mixed with treated steel for strength. The combination
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 414
Authors: Zachary Richardson
Co-Author: Leonard Jackson
permits the belt to be very thin (the CAD drawing of the conveyor belt is not to scale (though
everything else in the model is). The collection bin is made of the Carbon-Fiber and is hinged
allowing it and the conveyor belts to be folded in during the Supply Transfer Vehicles transfer
period.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 415
Authors: Zachary Richardson
Co-Author: Leonard Jackson
References
[1] [fomo.com/resources/technical-bulletins/openvsclosed.aspx. Accessed Feb. 2011.]
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 416
Authors: Leonard Jackson
5.6.6 ISPP Tank and Vehicle Connections
Since there will be recycling of the STV tanks and second stage Ares V tanks, our tanks
designed in the corresponding appendix section are not used. We reused these tanks to cut down
on the amount of launches for our mission. Refer to Section 5.2.7.1 for more information.
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 417
Author: Sarah Jo DeFini
5.6.7. ISPP Communication Systems
The ISPP Station has two pieces of communication hardware. The first is a wireless antenna
to communicate with the harvesters. Its specifications are the same as the antennas on the
harvesters, given in table 5.6.7-1.
Table 5.6.7-1: Antenna Specifications
Length, m 0.0485
Radius, m 0.005
Mass, kg 0.034
Efficiency 0.9997
Gain, dB 1.631
Beamwidth,
degrees 124
Radiated Power,
kW 0.0409
The second piece of hardware is a dish that communicates with the halo satellites and also
receives status signals from the storage tanks (which are each equipped with a 10 cm diameter
transmitter dish). The physical specifications for the ISPP dish are listed in Table 5.6.7-2
Table 5.6.7.5-2: Communications Dish Physical Parameters
STV Communications Dish
Diameter (m) 1
Mass (kg) 1.69
Power (kW) 5.0
Volume (m3) 0.005
Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 418
Author: Matthew Hill
5.6.8. ISPP End of Life Configuration
The ISPP stations meet their demise much the same way they began their life, by just sitting
there. After the Crew Transfer Vehicle departs and the Rescue Rover crashes into the planet, the
ISPP stations will continue to sit as silent monuments to Project Vision. The reactors shut down
after running out of fuel proceeded by the jettisoning of any excess propellant from the tanks to
avoid any dangerous explosions. Some components of the stations, such as the tanks, may be of
possible use to future missions but the need and possibility for such use must be analyzed further
in future projects.
Detailed Vehicle Descriptions Exploration Rovers Page 419
Author: Ben Stirgwolt
5.7 Exploration Rovers
5.7.1 Configuration
The Exploration Rovers are designed so that 2 crew members can live in it for up to 7 days at
a time comfortably. In order to maintain a sense of separation of work and non-work, the layout
of the exploration rovers, which we have named Castor and Pollux, is divided such that there are
separate areas for navigation, experimentation, meal preparation, and sleeping. There is also a
lavatory onboard. An example of the layout of the Exploration Rovers is in the appendix.
The sleeping quarters in located beneath the navigation area so as to minimize the overall
length of the vehicle. This area is accessible through a hatch located just behind the navigation
chairs. An artistic rendering of the interior of the navigation area is presented in Fig. 5.7.1-1,
where the hatch to the sleeping quarters is just behind the central control consol
Detailed Vehicle Descriptions Exploration Rovers Page 420
Author: Ben Stirgwolt
Figure 5.7.1-1 A crew member uses the robotic arm to pick up a rock during a mission on
the exploration rover
There are also two docking ports on the exploration rover, one port and one starboard in case
there is a problem with one of the doors. It also makes docking easier so that the Rover can
approach the CTV or another rover when docking from either side.
By: Ben Stirgwolt
Detailed Vehicle Descriptions Crew Capsule Page 421
Author: Joel Lau, David Wyant
5.7.2. Exploration Rover Power
To provide power to the Exploration Rover, we equip it with a 40.00 kW internal combustion
engine burning LH2 and LOX. The system produces mechanical power that we use to power the
wheels of the Exploration Rover. It also outputs excess mechanical power to an electric
generator which converts the excess mechanical power to electrical power for the electrical
loads. A mass and volume summary of the Exploration Rover’s power solution is shown below.
Table 5.7.2-1 Exploration Rover Mass and Volume Specifications
System IMLEO, kg Wet Mass, kg Volume, m3
Internal Combustion Engine [1] 98.00 98.00 0.36
Electric Generator 50.00 50.00 0.17
Na-S Batteries [2] 41.82 41.82 0.09
H2 Fuel 3.00 33.93 - -
O2 Oxidizer 24.00 271.44 - -
H2 Tank 5.55 5.55 0.45
O2 Tank 1.56 1.56 0.21
Total 223.93 502.30 1.28
We direct 16kW of mechanical power to the wheels of the Exploration Rover. This is easily
enough to power the rover at speeds over 4 m/s up a 45 degree incline. The power requirement
for the propulsion system is a function of vehicle weight, nominal operating speed, and the
maximum incline that can be traversed. In order to accommodate the required travel distance, a
nominal operating speed of 4 m/s and the ability to climb a 45 degree incline are the design
parameters. This leads to a minimum power requirement of 12.39 kW of mechanical power.
Additionally, we use an electric generator to convert 24 kWm to 12 kWe to supply power to
the electrical loads. The generator has an efficiency of 50%, a mass of 50 kg and a volume of
0.17 m3.A schematic of the electrical power system is shown in Fig. 5.7-1 below.
Detailed Vehicle Descriptions Crew Capsule Page 422
Author: Joel Lau
Figure 5.7.2-1 A generator converts mechanical power to electrical to satisfy the
Exploration Rovers’ electrical power needs
Detailed Vehicle Descriptions Crew Capsule Page 423
Author: Joel Lau
As shown above, the power system also includes three batteries. Each battery is a 2092 W-hr
Sodium-Sulfur (Na-S) battery, for a total energy storage of 6276 W-hrs. This is enough energy to
start the internal combustion engine and provide life support and limited communication for over
eight hours while the astronauts are sleeping. We equip the rovers with three separate batteries so
that if one fails, the other two are enough to maintain life support and some communication,
giving the rover pilots time to call for a rescue or travel back to the ISPP station. While the
engine is running during the day, it recharges the batteries so that the engine can be turned off
every night when the pilots sleep. These batteries have a total mass of 41.82 kg and a volume of
0.09m3.
Finally, we equip the electrical power system with a Peak Power Tracker. This tracker places
a load on the generator, and hence the engine, equal to that required by the loads. We do this to
ensure that the engine runs at the minimum speed necessary and does not burn excess fuel.
We show below the power requirements of the two exploration rovers.
Detailed Vehicle Descriptions Crew Capsule Page 424
Author: Joel Lau
Table 5.7.2-2 Exploration Power Requirements
System Power Requirement, kW Type
Communication
Transmitter Dish (RF) 0.10
Cellphones 0.28
2 Monitors 0.20
Total Communication 0.58 Electric
ADCS
Sensors (MRU's) 0.02
Computer system 0.01
Total ADCS 0.03 Electric
Propulsion
Drive Power 12.39 Mechanical
Structures
Thermal Control 0.03
Scissor Lift 1.07
Total Structures 1.10 Electric
Human Factors / Science
Environmental & Life Support 0.42
Science Equipment 0.68
Interior 2.30
Exterior 6.10
Total Human Factors / Science 9.50 Electric
Total Electric 11.21
Total Mechanical 12.39
Detailed Vehicle Descriptions Crew Capsule Page 425
Author: Joel Lau
References
[1] Ferguson, Colin R. "Internal Combustion Engines, Applied Thermosciences", New York:
Wiley, 1986.
[2] Wertz, James Richard, and Larson, Wiley J. "Space Mission Analysis
and Design", Microcosm, 1999
Detailed Vehicle Descriptions Exploration Rovers Page 426
Author: David Wyant
5.7.3 Propulsion Systems
The propulsion system that drives our Rovers is unique to the others presented in this mission.
It does not use any rocket propulsion; instead we implement an internal combustion engine to
provide both electrical generation and the mechanical drive power. In addition to the engine, the
drive system is comprised of all of the components of the suspension system and steering
mechanisms for the purposes of this report.
The overall mass of the propulsion system as well as its volume and power requirements are
seen in the following table.
Table 5.7.3-1 Propulsion system masses and volumes
Mass, kg External Volume, m3
Engine 98.00 - -
Transmission 73.65 - -
Chassis 1948.2 0.9093
Suspension
Wheels
200.0
81.92
- -
0.7226
Total 2303.77 1.63
Each of these components, the engine, transmission, chassis, suspension, wheels and how they
work together to create our explorations Rovers will be discussed in detail in the following
sections. Details concerning the performance parameters that drove the sizing of each
component can be found in the appendix.
Detailed Vehicle Descriptions Exploration Rovers Page 427
Author: David Wyant
5.7.3.1 Engine
We will syphon off our mechanical drive power from the internal combustion motor. In doing
a top level analysis, we find that converting all of the energy from the generator to electricity and
then creating a drive system based on electric power is wasteful. More mechanical power is
required to create electricity than to simply utilize the rotational speed of a drive shaft for
powering the drive systems. To this end, the required mechanical drive power creation will
come from the same engine as used for power generation.
The Exploration Rovers will be using 12.4 kW of mechanical power from this engine. This
will accommodate a speed of 4 m/s up an incline.
Detailed Vehicle Descriptions Exploration Rovers Page 428
Author: David Wyant
5.7.3.2 Transmission
In order to provide a steering method as well as an ability to vary the power provided to the
wheels, we will install four independently operating hydrostatic transmissions with a total mass
of approximately 74 kg.
Hydrostatic transmissions provide us with an infinitely variable gear ratio. This effectively
allows for throttling the power to each wheel independently. This allows for steering abilities
akin to that of a skid loader or other similar craft with a turning radius of zero degrees. In a
closed hydrostatic transmission, the torque can be transmitted both forward and in reverse,
eliminating the need to design a braking system of our vehicle [1].
To use hydrostatic transmission on Ceres, we will need to bring hydraulic fluids that are not
susceptible to the extreme cold and other elements in space with us. This could be accomplished
through the use of high molecular weight polyalphaolefin (PAO) or multiply alkylated
cyclopentanes (MACs). PAOs and MACS are both currently being used in spacecraft as oil and
lubricants making them viable options for our hydraulic fluids [2].
We will have no issues scaling these devices to deliver the proper torque and size for this
vehicle as they are used in all types of vehicles from lawn mowers to modern combine harvesters
[3].
Detailed Vehicle Descriptions Exploration Rovers Page 429
Author: David Wyant
5.7.3.3 Chassis
Our Rover design will include a chassis as a central point to affix all external, physical Rover
components. The body, suspension, wheels and axles all attach to this frame. Essentially a
rectangular frame encompassing the same dimensions as the body of the Rover, it has a large
mass of 1948 kg.
The mass of this chassis was calculated using the results of Donald Malen’s study of vehicle
component mass percentages [4]. This estimation method will be discussed more in depth in the
appendix.
Detailed Vehicle Descriptions Exploration Rovers Page 430
Author: David Wyant
5.7.3.4 Suspension
The overall effectiveness of the suspension system can be characterized by the coefficient of
restitution. This coefficient measures the ratio of speed of separation to speed of approach in a
collision [5] or “bounciness” of the object. In order to keep our Rover from bouncing too much,
we designed a suspension system with the goal of keeping the effective coefficient of restitution
between 0 and 0.2. This will be different for each collision; however, we will assume that an
average value between the two is attainable.
The mass of the suspension system was sized using the method as the chassis. These
calculations provide a final suspension system mass of 200 kg.
Detailed Vehicle Descriptions Exploration Rovers Page 431
Author: David Wyant
5.7.3.5 Wheels
The wheels used by our Exploration Rovers are based on the wheels of the lunar rover from
the Apollo missions. They will be a hollow wheel made of metal with mesh sides. The tread of
the wheel will have chevrons cut out of it to help it dig into the regolith and gain additional
traction.
Figure 5.7.3-1 This is a side and front view image of the Rover wheels. The darkened
wedges in the right hand figure illustrate the chevrons, where metal has been removed.
The design specifications of these wheels can be seen in the image above, the dimensions are
in meters. The mass of each wheel is 20.48 kg for a total mass of 81.92 kg for all of the wheels
on the Rovers. Calculations will be listed in the appendix.
Detailed Vehicle Descriptions Exploration Rovers Page 432
Author: David Wyant
References
[1] Rydberg, K. E., "Hydrostatic Drives in Heavy Mobile Machinery – New Concepts and
Development Trends," Society of Automotive Engineers, Inc. Paper 98-1989, 1997.
[2] Bhushan, B., Modern Tribology Handbook, Volume 1, CRC Press LLC, 2001.
[3] "John Deere 9870-STS," Products and Equipment, Deere and Company, Illinois, 2011.
[http://www.deere.com/servlet/ProdCatProduct?tM=FR&pNbr=9870SH. Accessed
3/31/11.]
[4] Malen, D., "Preliminary Vehicle Mass Estimation Using Empirical Subsystem Influence
Coefficients," Auto/Steel Partnership, May 2007.
[5] "Coefficient of Restitution," Eric Weisstein’s World of Physics,
[http://scienceworld.wolfram.com/physics/CoefficientofRestitution.html. Accessed
2/31/11.]
Detailed Vehicle Descriptions Exploration Rovers Page 433
Author: Ben Stirgwolt
Co-Author: Jillian Roberts
5.7.4 Human Factors Systems and Habitability Considerations
The Exploration Rovers are designed to be a comfortable environment for the astronauts;
however, there are limitations in order to keep the rover to a reasonable size. The rovers are
stocked with enough food for 7 days, but there is only a microwave oven onboard, so there is no
extensive cooking that takes place during the missions. There is also no shower onboard due to
size limitations. There is enough storage area for the crew to bring several days’ worth of
clothing in addition to any personal items they want.
We use the same air ventilation system onboard the CTV in order to ensure sufficient airflow
in the sleeping quarters, which is in a small, confined area. The ventilation fans are small and
quiet so as to generate as little noise as possible.
Because the astronauts could spend a full week in the Exploration Rovers, we must make
water provisions to allow for drinking, food rehydration, and basic hygiene while away from the
CTV. We conducted a trade study which determined that recycling water stores, instead of
storing it all, would significantly decrease mass in the rover. This trade study can be found in the
Appendix. The total mass, volume, and power requirements for the Exploration Rovers are
found in the table below. This data is for water aboard each rover.
Table 5.7.4-1 Specifications for the water supply and recycling system on the Exploration
Rovers
Crew
Members
Days Mass, kg Power, kW Volume, m3
Water Supply and
Regeneration
4 7 408.1 0.160 0.608
Detailed Vehicle Descriptions Exploration Rovers Page 434
Author: Ben Stirgwolt
Co-Author: Jillian Roberts
In case of a fire, the Exploration Rovers each have two fire extinguishers and one smoke
detector. See the Fire Suppression and Detection section from the Crew Transfer Vehicle and its
corresponding Appendix for details. The mass, power, and volume can be found in the table
below.
Table 5.4.7-2 Specifications for the fire suppression and detection system
Mass, kg Power, kW Volume, m3
Fire Detection and Suppression 23.27 0.0015 0.0788
To provide an ergonomic working environment which is well-lit, the Exploration Rovers will
have a lighting system. The table below describes the mass, power, and volume of the lighting
system. We assume the Rover needs 1000 lux for the astronauts to efficiently perform science
experiments.
Table 5.4.7-3 Specifications for the lighting system
Mass, kg Power, kW Volume, m3
Lighting System 183.6 1.530 5.100
Detailed Vehicle Descriptions Exploration Rovers Page 435
Author: David Schafer
5.7.5 Attitude Determination and Control Systems
The Exploration Rovers need to track their position throughout their lifecycles. Knowledge
of position is only fed to control an antenna to point at the satellites, so the exploration rovers
simply require attitude determination sensors. An additional need for an autonomous system for
one specific maneuver is also required of the system, so the attitude determination system must
be able to track the Rover’s position throughout the mission, manned or unmanned. The Rovers
have no need for a specific actuating system, as these have been provided in the basic propulsion
system of the vehicles.
Detailed Vehicle Descriptions Exploration Rovers Page 436
Author: David Schafer
5.7.5.1 Attitude Determination
Each rover will implement a system of inertial (motion) reference units (MRU) coupled with
computers in order to determine both the attitude and position of the rovers. Use of the inertial-
style attitude determination systems (accelerometers) requires computers to integrate the
information to find position and velocity, and is both highly reliable and also fairly resistant to
damage. These systems are mounted in the storage area beneath the floor of the
Rovers, and oriented perpendicularly (along the axis of symmetry) to each other in order to
maintain the high accuracy required in all directions for the attitude of the craft. This way, the
motion reference units can maintain an accuracy of about 0.02° while the craft needs only to
keep accuracy to about 0.2°, so we create the system with plenty of room for error. Using the
specific Kongsberg model 5+ MRUs as an example (these could be used, but would necessitate
altering the internal gains of the system to account for the changes in gravity from Earth to
Ceres), the attitude determination system produces the values shown below in table 5.7.5.1-1.
Table 5.7.5.1-1 Exploration Rover attitude determination system
Hardware Mass, kg Power, kW Volume, m3
MRU’s 5 0.024 0.02
Computers 3 0.01 0.004
Total 8 0.034 0.024
Detailed Vehicle Descriptions Exploration Rovers Page 437
Author: David Schafer
5.7.5.2 Attitude Control
As we drive the Exploration Rovers across the surface of Ceres, the attitude control aspect is
very similar to that of a standard car. The vehicle only has two degrees of freedom in translation
and two degrees of freedom in orientation. By adjusting the directions of the wheels and the
power of the engines, the rovers can traverse on a majority of the surface of Ceres. These
systems are set up so that the crew can operate the rovers, or an autonomous system can drive the
crew.
Detailed Vehicle Descriptions Exploration Rovers Page 438
Author: David Schafer
5.7.5.3 Autonomous Control Considerations
In order to implement the unmanned transfer from Station 1 to Station 2 at the mid-duration
point of our mission, we must integrate the attitude determination system with an autonomous
controller. This controller must read both the position and attitude, as well as the velocity and
acceleration of the Rover. All of the necessary information should already be provided by the
sensors on the rovers and already input into the system computers. Additionally, the autonomous
system must analyze additional information from the rover engines, the communications dish,
and the power system of the vehicle, which will all need to be combined on the system
computers and fed to the wheel motors for the system to operate properly. This is covered in
greater depth in section 5.7.10.
Detailed Vehicle Descriptions Exploration Rovers Page 439
Author: Kim Madden
5.7.6 Structural and Thermal Systems
5.7.6.1 Structural Components
Our Exploration Rover consists of a pressure vessel main body, two clear ellipsoidal
windshields, a floor to divide living area from storage area, two rock boxes for containing Ceres
rock samples, and radiation shielding. All of these components come together to create the
general cylindrical shape of the Rover. The structure of the Exploration Rover safely contains the
crew and essential life support systems.
Pressure Vessel
Pressure vessels are traditionally circular because of the excess stresses introduced by
bending. We choose a cylinder shape for our rovers for this reason, as well as to obtain the most
usable space inside. A sphere is the ideal shape for pressure vessels, but needs a very large radius
in order to contain the required equipment. With a cylinder, the radius can be smaller since the
length can be changed.
First, we need to determine the radius required for the cylinder. Based on the inside
configuration, we need a floor length of 4 m. We decide to place a floor along a chord of the
cross section, instead of across the middle. Having the floor lower makes the necessary radius
smaller and gives more head-room for a crew member to comfortably stand. Figure 5.7.6.1-1
shows a circular cross section with the approximate floor location.
Detailed Vehicle Descriptions Exploration Rovers Page 440
Author: Kim Madden
Figure 5.7.6.1-1 A circular cross section, with important variables defined. We wish to find
the radius using the desired storage height h and the chord length c.
We choose to have a storage height h of 1.5 m. This creates enough storage space for the life
support systems, and minimizes the actual radius of the cylinder. While calculating the radius,
we also ensure that a person is able to stand up comfortably (called the head room).
The radius of the cylinder is 2.137 m. With the storage height of 1.5 m, there is 2.77 m of
head room above the floor, which is enough for a man to stand up. The sides have a smaller
height, but there will be counters and dock doors there, so this reduced clearance will not be an
issue. The length of the cylinder is 3 m. This is dictated by the previously designed floor plan.
We make the thickness of the walls to be 1.5 cm. We can now determine the mass of the
pressure vessel part of the Rover by multiplying this calculated thickness by the surface area.
The mass of the pressure vessel is 1,698.02 kg. The internal volume is 43.05 m3.
Windshields
We add windshields to the front and back of the Rovers to serve a number of different
purposes. The first reason is that we want the crew to be able to see where they are driving. This
way, they can avoid rocks or other obstacles, as well as have a good view of the Ceres surface.
Another reason to include them is to keep Ceres dust and dirt out of the Rover, maintaining a
R
C
h
D
Diagram by Kim Madden
Detailed Vehicle Descriptions Exploration Rovers Page 441
Author: Kim Madden
clean and safe environment for our crew. Lastly, these end caps for the Rover will create a closed
area to maintain the internal pressure.
We choose to make our windshields out of polycarbonate, which is a stronger and more
durable material than plexiglass. The material needs to be clear so that the crew can see out of it.
Polycarbonate has a yield strength of 62.1 MPa, and a density of 1200 kg/m3 [1].
The shape we choose for the windshields is a 2:1 ellipsoidal head, which can be seen in Fig.
5.7.6.1-3. This design would have the same benefits as the hemispherical shape, but would
reduce extra mass because it does not stick out as far.
Figure 5.7.6.1-3 Diagram of the 2:1 ellipsoidal windshield. It reduces mass while
maintaining the visibility.
We design the ellipsoidal to have a thickness of 1.5 cm. The surface area of two ellipsoidal
windshields is 24.53 m2. By multiplying the surface area by the required thickness and the
density of polycarbonate, we determine the mass of the two ellipsoidal windshields. The mass is
882.93 kg, and the internal volume is 10.22 m3.
Floors
We must include floors so that the crew can stand and work on them. As previously
discussed, we already know the dimensions of the floor, which are based on the configuration of
the Rover’s interior. The floor is a rectangle 1.5 m above the bottom of the cylinder. In order to
determine the thickness of the floor, we model it as a beam fixed at both ends with a
2R R/2 R/2
Diagram by Kim Madden
Detailed Vehicle Descriptions Exploration Rovers Page 442
Author: Kim Madden
concentrated load in the middle. This gives a very conservative estimate of the floor thickness,
because all of the mass will not be concentrated in the middle, but rather, spread around. We
assume that 2/3 of the human factors and science mass is on the floor, while 1/3 of it is in
storage. We also assume that the maximum deflection of the floor can be 1 cm. This deflection
occurs in the center [2].
We build the floor to be 2 cm thick, which results in a floor mass of 997.83 kg, and a volume
of 0.07 m3.
Storage and Sleeping Dividing Wall
We include a dividing wall between the storage area and the bed area in the bottom of the
Exploration Rover. This is for the comfort of the crew, so they are not sleeping next to life
support systems. The area of that wall is 4.49 m2. We make the wall 1 cm thick, as it has not
supporting any structure, and out of aluminum. The mass of this dividing wall is 126.23 kg.
Radiation Shielding
Because we are working in a vacuum, we need some radiation shielding to protect the crew
and electronics from harmful exposure. There are many different ideas for the ideal radiation
protection, but it is hard to get an exact value since studies can only be done in space. A heavily
shielded area is needed in case of a solar particle event and galactic cosmic rays. This is located
in the CTV; the Rovers will not require this much shielding.
We choose to use a passive shield consisting of aluminum and polyethylene. Polyethylene
contains a lot of hydrogen and is lightweight, making an excellent shield for radiation. This
material has a density of 925 kg/m3
[1]. The outer layer of the Rover is already made of 1.5 cm
of aluminum, so this also doubles as radiation shielding. However, we need a thicker shield to be
effective. For a light shielding, we use 80 kg/m2 of material [3,4]. The aluminum pressure vessel
Detailed Vehicle Descriptions Exploration Rovers Page 443
Author: Kim Madden
is 42.15 kg/m2, so we require an additional 37.85 kg/m
2. To add the additional mass, we need
another 4 cm of polyethylene, which is located inside the pressure vessel portion of the Rover
and has a mass of 1,524.8 kg.
Buckling
Now that we have the basic structure of the Rover complete, we need to make sure that it
doesn’t buckle during the launch. We assume that during the STV launch, it will experience 6g’s
of acceleration. The force during launch is the total mass of the Rover multiplied by the
acceleration during launch, which is 502.5 kN.
The force that would buckle the Rover during launch is 9,138,238 kN. Since this force is
MUCH larger than the force the Rover will experience during launch, we conclude that the
Rover will not buckle during the launch.
Rock Storage Boxes
The rock storage boxes are required to hold the collected Ceres surface samples. The robotic
arm in front and back of the Exploration Rover will grab rocks and place them into these storage
containers.
The best way to fix the boxes onto the front and back of the Rover is to place them under the
windshield. This puts a constraint on the height, which is 0.35 m. We make the thickness of the
box 2 cm, which should be thick enough to contain the rocks, especially with Ceres’ low gravity.
The total mass of two boxes is 433.62 kg, and the combined volume of the two boxes is 1.35 m3.
Nuts, Bolts and Screws
In order to account for various building materials, such as nuts, bolts and screws, we add 10%
of the total structure mass to the totals, as well as 5% of the structural volume. These are
Detailed Vehicle Descriptions Exploration Rovers Page 444
Author: Kim Madden
approximated values, and while they may seem negligible, they actually add up to 553.72 kg and
0.23 m3.
Structural Summary
Table 5.7.6.1-1 shows a summary of the mass, power, and volume requirements of the
structural components of the Exploration Rover. Figure 5.7.6.1-4 shows a picture of the
Exploration Rover with the structural components pointed out.
Table 5.7.6.1-1 Structural summary of mass, power and volume
Component Mass, kg Power, kW Volume, m3
Pressure Vessel 1,698.02 0 43.05
Windshields 882.94 0 10.22
Floors 997.83 0 0.06
Dividing Wall 126.23 0 0.05
Radiation Shielding 1,524.80 0 1.65
Rock Storage Boxes 433.62 0 1.35
Nuts, Bolts and Screws 553.72 0 0.23
Totals 6,090.93 0 56.56
Detailed Vehicle Descriptions Exploration Rovers Page 445
Author: Kim Madden
Figure 5.7.6.1-4 Model of the Exploration Rover, with the structural components pointed
out. All of the other components are on the inside of the Rover
Windshield
Rock Storage Box
Pressure Vessel
Detailed Vehicle Descriptions Exploration Rovers Page 446
Author: Kim Madden
References
[1] Callister, W. D., Materials Science and Engineering An Introduction, 7th
ed., John Wiley &
Sons, Inc., Pennsylvania, 2007, Appendix B.
[2] Gere, J. M. and Goodno, B. J., Mechanics of Materials, 7th
ed., Cengage Learning, Ontario,
[3] Wilson, J. W., Miller, J., Konradi, A., and Cucinotta, A. F., “Shielding Strategies for
Human Space Exploration”, National Aeronautics and Space Exploration, December, 1997.
[4] National Council on Radiation Protection and Measurements, NCRP Report No. 98:
“Guidance On Radiation Received In Space Activities”, Bethesda, MD: NCRP, 1989.
Detailed Vehicle Descriptions Exploration Rovers Page 447
Author: Kim Madden
5.7.6.2 Thermal Control System
A thermal control system is important for all space vehicles, especially manned vehicles.
While space is quite cold, there are electronics and motors inside each vehicle that produce heat.
We want to design a thermal control system that keeps the internal temperature comfortable for
the crew.
There are two main sources of power that add heat to the inside of the vehicle. The first is
the heat that is produced from the crew inside, which is 61.3 watts of heat per person. We
multiply this number by the number of people inside the vehicle at any time to get the amount of
heat that needs to be rejected. For the Exploration Rover, there are a maximum of 4 crew
members inside at any time, so 245.2 watts of heat need to be rejected. Power also comes from
the rejected heat from the electronics, which is produced because the electronics are not 100%
efficient. We assume that the electronics are 65% efficient, as advised by Dr. Boris Yendler. The
electronics in the Exploration Rover require 3.609 kW of power, so 1.349 kW need to be
rejected.
There are two ways that heat leaves the Exploration Rover. The first is due to the colder
temperatures on Ceres. The temperature on Ceres ranges 235 K during the day and 100 K at
night. We want to maintain the inside of the Rover at a comfortable temperature for the crew.
We choose to keep it at 293 K, which is a comfortable room temperature on Earth. Heat will
escape the Rover because of the difference in temperatures. We require an additional heat
rejection system, and that is radiators and heat pumps. Heat pipes carry heat from the electronics
to radiators on top of the Rover, which then reject excess heat. Figure 5.7.6.2-1 shows a
schematic of the thermal control systems.
Detailed Vehicle Descriptions Exploration Rovers Page 448
Author: Kim Madden
Figure 5.7.6.2-1 This is a schematic of the thermal control systems. The top portion shows
what kind of heat transfer goes into and out of the vehicle, and the bottom portion shows
the inside system.
Multilayer Insulation
We want to minimize the amount of heat lost through the Rover due to environmental
differences. We wrap the Rover in multilayer insulation (MLI) in order to stop some of this heat
flow. MLI blankets are 30 layers of 0.25 mm thick metalized Mylar sheets separated by a mesh.
This acts as a barrier for the heat radiated from the surface of the spacecraft into the cold space.
The outer layer is thicker since it will be exposed to the elements, and white to reflect sunlight
[1]. In order to determine the mass of the MLI covering the Rover, we multiply the surface area
of the Rover that will be covered in MLI (40.285 m2) by the density of MLI.
Heat Pipes
Heat pipes will run all through the Exploration Rover in order to carry heat from the
electronics to the radiators. We choose to use water as the working fluid for the heat pipes.
Diagram by Kim Madden
Detailed Vehicle Descriptions Exploration Rovers Page 449
Author: Kim Madden
Ammonia is the traditional working fluid; however, if there was a leak, the crew would be in
trouble. Water will be available on the surface, so if it needs to be replenished, it can be easily.
Also, a water leak will not harm the crew in any way.
When the water flows under an electronic, it will heat up and vaporize. As the water moves
away from the electronics towards the radiators, the water will condense. This is how heat moves
throughout the pipes. Small resistance heaters are located near the radiators to keep the water
liquid when it starts to get colder. If the water freezes when exposed to the radiators, the heat
pipe would then be useless and the Rover will overheat. The mass of the heat pipe, including the
water required, is 29.876 kg.
Radiators
We must now determine the mass and size of the radiators. The radiators are located on top of
the Rovers. The radiators will be required to open and close depending on how much heat needs
to be rejected. Rubber corners connect the heat pipes through the radiators to the heat pipes in the
vehicle. This allows the radiators to fold, and also stops the flow of water when the radiators are
folded. They can be closed during the night to stop heat flow to keep the inside warm, and then
open up again during the day. On the Exploration Rover, there will be 4 sets of 2 radiators (8
total radiator panels). One side of each radiator set will be covered in MLI to stop more heat
flow. This leaves ¾ sides of each radiator set to radiate heat. The power requirements for this
mechanism can be found in section G.3.2 by Joel Lau. These radiators require 0.061 kW of
power to raise them.
To remove a certain amount of heat, the radiator needs to have a certain surface area. The
size of one radiator panel is 0.346 m by 0.346 m, giving a surface area of 0.956 m2. The mass of
Detailed Vehicle Descriptions Exploration Rovers Page 450
Author: Kim Madden
all the radiator panels is 304.034 kg. This includes the MLI covering ¼ of the surface area of a
radiator set.
Aluminum Plates
For heat to be transferred to the heat pipe from the electronics, an aluminum plate needs to be
underneath. We assume that there is 0.5 square meter of aluminum throughout the Rover. This is
broken up and placed under every electronic, with the heat pipes flowing under the plate. The
thickness of the plate is 5 mm.
Heater
We also include a heater in the Exploration Rover to add heat to the inside of the vehicle in
case it gets too cold for the crew. We have created a simple system to accomplish this. We
develop a system that is slightly opposite of the heat removal process. The combustion engine
that gives power to the Rover has a low efficiency, and thus puts out a lot of heat. We run a heat
pipe through the engine to gather this extra heat and deliver it to the Rover, similar to a car
exhaust. A small radiator panel inside the vehicle is able to be manually lifted to let heat in, and
closed when the temperature is comfortable for the crew. It is covered in MLI so that heat is not
added to the vehicle when is it not wanted. Figure 5.7.6.2-2 shows a schematic of the heater.
Detailed Vehicle Descriptions Exploration Rovers Page 451
Author: Kim Madden
Figure 5.7.6.2-2 This is a schematic of the heater for the Exploration Rover. Heat is
transferred via heat pipe from the internal combustion engine to the inside of the Rover.
We design the heat pipe to collect 100 W of heat from the combustion engine. We choose
the length of this heat pipe to be 10 m, as it does not need to be snaked around the vehicle. The
total mass of the heater system is 15.378 kg.
Diagram by Kim Madden
Detailed Vehicle Descriptions Exploration Rovers Page 452
Author: Kim Madden
Results and Summary
Table 5.7.6.2-1 shows a compiled chart of the mass, power, and volume requirements for the
Exploration Rover thermal control system.
Table 5.7.6.2-1 Exploration Rover Thermal Control System Summary
Component Mass, kg Power, kW Volume, m3
MLI Covering 11.240 0 0.041
Heat Pipe 29.876 0 0.302
Radiators 304.038 0.061 0.108
Aluminum Plates 7.025 0 0.003
Heater 15.378 0 0.028
TOTAL 367.557 0.061 0.481
Detailed Vehicle Descriptions Exploration Rovers Page 453
Author: Kim Madden
References
[1] Birur, G. C., Siebes, G, and Swanson, T. D., “Spacecraft Thermal Control”, Encyclopedia
of Physical Science and Techonology, 3rd
ed., Academic Press, 30 March 2001.
Detailed Vehicle Descriptions Exploration Rovers Page 454
Author: Tony D’Mello
5.7.7 Communication Systems
5.7.7.1 Internal Communication
In table 5.7.7.1-1, we show the communication devices that we provide in each exploration
rover. Although each rover can accommodate four crew members, except under special
circumstances, only two crew members are present in each rover. Therefore, we only require
two televisions in each rover. Each rover has a small antenna located on the ceiling in the center
of the rover that can accommodate communication among four cell phone devices.
Table 5.7.7.1-1 Internal communication device characteristics for one Rover
2 Televisions 4 Cell Phones 1 Antenna
Mass, kg 10 2.4 1.7e-3
Power, kW 0.2 0.28 1.5e-3
Volume, m3
0.016 0.002 1.94e-7
Detailed Vehicle Descriptions Exploration Rovers Page 455
Author: Tony D’Mello
5.7.7.2. External Communication
Although both rovers send six high definition television (HDTV) signals, Rover 1 operates at
higher frequencies than Rover 2 so that the two signals do not interfere with each other. We can
designate four of the channels for crew communication while the other two refer to the cameras
located outside of the rovers. The crew can also operate both camera feeds when remotely
controlling the rovers from the Crew Transfer Vehicle (CTV). Table 5.7.7.2-1 displays the mass,
power, and volume values for the transmitter and the receiver.
Table 5.7.7.2-1 External communication device characteristics for Rover 1
1 Transmitter 1 Receiver
Frequency, GHz 40 11
Data Rate, HDTV channels 6 4
Mass, kg 0.33 6.35
Power, kW 0.1 __
Dish Diameter, m 0.54 1.25
As mentioned earlier, Rover 2 operates at lower frequencies than Rover 1 which causes an
increase in mass and volume as can be seen in Table 5.7.7.2-2.
Detailed Vehicle Descriptions Exploration Rovers Page 456
Author: Tony D’Mello
Table 5.7.7.2-2 External communication device characteristics for Rover 2
1 Transmitter 1 Receiver
Frequency, GHz 26.5 7
Data Rate, HDTV channels 6 4
Mass, kg 0.67 16.35
Power, kW 0.1 __
Dish Diameter, m 0.63 1.6
Detailed Vehicle Descriptions Exploration Rovers Page 457
Author: Ben Stirgwolt
5.7.8 Ceres Rock Collection Process
Upon returning to the CTV after a sortie, the crew members use the robotic arms to lift the
rock storage containers from the fore and aft of the exploration rover. The rocks are dumped into
a pile near the CTV and the storage contain is then returned to its position on the rover. For each
sortie, the rocks are placed into individual piles so as to identify the location of where the
regolith originated. Once the time at ISPP station 1 is nearing completion, the crew members
must use the robotic arms to sort through each individual pile, looking for the most valuable
rocks that should be returned to Earth for further inspection. Half a ton of rocks is selected from
ISPP station 1. We then place this half-ton into a cryogenic storage container that is accessible
from the crew capsule. The process is then repeated at ISPP station 2.
Detailed Vehicle Descriptions Exploration Rovers Page 458
Author: Ben Stirgwolt
5.7.9 Science Toolbox and Experimentation
The rovers deploy several experiments to study the composition of the regolith. A meteorite
experiment studies the small particles that strike the surface of Ceres, measuring their velocity
and direction at the time of impact. An electrical properties experiment uses transmitting
antennae to determine the electrical properties of the regolith. In addition to these experiments,
the rovers have the following geological tools onboard:
Heat flow probes
Electromagnetic sounder
Thermal emission spectrometer
Alpha Particle X-ray spectrometer
Microscope
Magnetic array
Rock abrasion tool
Panoramic cameras
For immediate scientific inquiry of the soil composition, there is a glove box in the
exploration rover so astronauts can examine rock samples in their natural environment, without
having to expose the soil to a foreign environment. The astronauts use the robotic arms to select
the rock of interest and then maneuver it to the glove box tray.
With regards to physics and astronomy experiments, the astronauts use a traverse gravimeter
that is deployed at several locations on Ceres to make relative gravity measurements. They also
use a small research telescope and an ultraviolet light telescope study the evolution of galaxies.
Detailed Vehicle Descriptions Exploration Rovers Page 459
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5.7.10 Autonomous Operations
In order for the crew to transfer from Station 1 to Station 2 at the midpoint of the mission, we
need the rovers to be able to traverse Ceres without any manned guidance. This means an
autonomous system is needed to guide the Rovers halfway across the surface of Ceres. This
system must be fully integrated into the positioning, power, propulsion, and even the
communications systems in order to operate the Rover on its solo journey. Additionally, the
Rovers need some way to comprehend and analyze the directions it can travel in to reach its
destination without any outside knowledge.
The Exploration Rovers traverse half of the circumference of Ceres (approximately 1530 km)
for their autonomous move from ISPP Station 1 to ISPP Station 2. If we assume that the Rovers
will be running at nominal operating speed, the total elapsed time for the transfer will be 4.43
days.
Detailed Vehicle Descriptions Exploration Rovers Page 460
Author: David Schafer
5.7.10.1 Land feature correction
No Ceres surface mapping has yet been done. This means that the Rovers will have to find
their own way across the dwarf planet with minimal input from human users. In order to do this,
we suggest implementing a visual comparison system to work with the cameras already onboard
the Rovers to determine the best course of action. This preliminary study did not delve deep
enough to look at actual logic processes to aid the Rovers in both determining the ground around
it nor in deciding the best direction to take, yet these types of systems have been in development
for years. A specific instance is the driverless car concept heavily researched since 1995. As
these systems are still far from well developed, they hold a Technology Readiness Level of about
4. This indicates that the concept has been proven feasible, but much progress is needed to truly
make a driverless car. The current major design problem is the automated decision making
process. While the Rover can certainly use the sensors and cameras to find out where it is, what
is ahead of it, and what is to the side of it, getting the Rover to decide which course of action is
best still remains the hardest challenge. Instances of small successes with decision making have
proven successful through the DARPA Urban Challenge [1].
Detailed Vehicle Descriptions Exploration Rovers Page 461
Author: David Schafer
5.7.10.2 Integration of all components
We set up the dual computer system for attitude determination on each of the Rovers to
combine the information from the attitude system and the land feature correction system. This
information is then put through a controller to find the necessary changes the propulsion system
and the communications system must make in order to both move to a pre-set waypoint and
continue communications with the crew. These values are then sent to the respective parts of the
rover, so the Rover can move, the antenna can track it, and the rover can reach a new position to
iterate through the entire procedure again. This is a standard automated control technique, with
the addition that the Rover must be able to determine its own easiest path to each waypoint.
Detailed Vehicle Descriptions Exploration Rovers Page 462
Author: David Schafer
References
[1] “DARPA Urban Challenge” DARPA Grand Challenge. 2007
Detailed Vehicle Descriptions Exploration Rovers Page 463
Author: Kim Madden
5.7.11 End of Life Configuration
Unfortunately, the exploration rovers do not have special end of life configurations. They will
stay on the Ceres surface forever. It is possible that they could be reused for future missions; they
would need to be restocked with supplies, refueled and repaired before any major expeditions
could take place.
There will be no hazards to the Ceres environment with leaving the rovers there, except the
extra space trash. There are no nuclear reactors to blow up, and any leftover fuel could be
expelled prior to leaving to avoid an explosion. All electronics inside will be shut off, and life
support systems will be removed to avoid the potential growth of mold. Ceres alien children
could safely use them as a playground.
Detailed Vehicle Descriptions Page 464
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Detailed Vehicle Descriptions Rescue Rover Page 465
Author: Ben Stirgwolt
5.8 Rescue Rover
5.8.1 Configuration
The Rescue Rover, like the Exploration Rovers, is designed so that there are separate areas for
navigation, travel, and medical operations. There are two medical beds in the back of the rover,
both located near a sink and all of the medical supplies. Two docking ports make it easy to dock
with both the CTV and the other rover when necessary. A example of a possible layout of the
Rescue Rover is in the appendix.
Detailed Vehicle Descriptions Rescue Rover Page 466
Author: Joel Lau
5.8.2. Rescue Rover Power
We supply power to the rescue rover with a combination of a hydrogen fuel cell and an array
of batteries. The fuel cell outputs 10 kW of electrical power. It is a proton exchange membrane
fuel cell and converts LH2 and LOX into water and electricity. The fuel cell consumes 0.019 kg
of LH2 and 0.140 kg of LOX per hour of operation [1]. At full capacity, the propellant tanks
hold enough LH2 and LOX for 48 hours of operation. Unlike the exploration rover, we provide
the rescue rover with propellant prior to LEO. This is so that the Rescue Rover will be
immediately ready for its first mission on Ceres. Shown below are the mass and volume
specifications of the Rescue Rover power system.
Table 5.8.2-1 Rescue Rover mass and volume specifications
System Wet Mass, kg Volume, m3
Hydrogen Fuel Cell 102.04 0.17
Na-S Batteries 41.82 0.09
H2 Fuel 0.92 -
O2 Oxidizer 6.71 -
H2 Tank 0.17 0.0014
O2 Tank 0.04 0.0061
H2O Tank 0.06 0.0077
Total 151.71 0.27
The power system also includes an array of three batteries. These store excess power and
supply backup power after the fuel cell runs out of fuel or in case of failure. Each battery is a
2092 W-hr Sodium-Sulfur (Na-S) battery, for total energy storage of 6276 W-hrs. The batteries
are capable of providing life support and limited communication for over eight hours if needed.
If one battery fails, the other two are enough to maintain life support and some communication,
giving the rover pilots time to organize a rescue or travel back to the ISPP station. The batteries
Detailed Vehicle Descriptions Rescue Rover Page 467
Author: Joel Lau, David Wyant
have a total mass of 41.82 kg and a volume of 0.09m3[2]. Shown in Fig. 5.8.2-1 is a schematic of
SPRINT’s power system.
Figure 5.8.2-1 Rescue Rover electrical power system
Shown below in Table 5.8.2-2 is the Rescue Rover’s power budget. As shown, the rescue
rover requires a total of 8.65 kW of electrical power. This requirement is more than met by the
10 kW fuel cell. The majority of the power load is consumed by the maneuvering motor (4.66
kW) and Human Factor and Science equipment (3.12 kW). Much like the exploration rovers, the
power requirement for the maneuvering motor must meet specific design parameters. In this
case, the conditions are a 45 degree incline at a nominal speed of 2 m/s. This leads to a
minimum power requirement of 4.66 kW. The rover will be able to supply this much power at
all times if necessary, allowing it to run at speeds above nominal over flat land for the purposes
of rescue missions.
Detailed Vehicle Descriptions Rescue Rover Page 468
Author: Joel Lau, David Wyant
Table 5.8.2-2 Rescue Rover power requirements
System Power Requirement, kW
Communication
Transmitter Dish (RF) 0.10
Cellphones 0.42
2 Monitors 0.20
Total Communication 0.72
ADCS
Sensors (MRU's) 0.02
Computer system 0.01
Total ADCS 0.03
Propulsion
Maneuvering Motor 4.66
Structures
Thermal Control 0.02
Human Factors / Science
Environmental & Life Support 0.42
Medical Equipment 0.30
Interior 1.30
Exterior 1.10
Total Human Factors / Science 3.12
Total Electric 8.65
Detailed Vehicle Descriptions Rescue Rover Page 469
Author: Joel Lau, David Wyant
Resources
[1] Wertz, James Richard, and Larson, Wiley J. "Space Mission Analysis
and Design", Microcosm, 1999.
[2] "NASA's Space Shuttle Orbiter", UTC Power, 2008, accessed 02 Apr 11 at
http://www.utcpower.com/fs/com/bin/fs_com_Page/0,11491,0115,00.html
Detailed Vehicle Descriptions Rescue Rover Page 470
Author: David Wyant
Co-Author: Evan Helmeid
5.8.3 Propulsion Systems
The Rescue Rover will have two propulsion systems onboard as well as an attitude and
control system. The two propulsion systems present on the craft are a main thruster to allow the
Rescue Rover to put itself into a powered flight and cover large areas of Ceres surface quickly
and a set of maneuvering motors to dock with the stranded rovers upon finding them. The
attitude and control thrusters will also be covered in this section of the paper.
Table 5.8.3-1 Propulsion System Masses and Volumes
Mass , kg External Volume, m3
Main Engine 12.15 0.1015
Attitude Thrusters 28.0 0.002
Maneuvering Motors
Wheels
2.88
81.92
- -
0.7226
Total 2303.77 0.8261
5.8.3.1 Main Thruster
A main thruster is used to launch the Rescue Rover either into orbit on onto its ballistic
trajectory. We size the engine of the Rescue Rover to provide sufficient thrust to weight ratio
(T:W) when heavy to allow proper vehicle control. The engine is throttleable to 10% of its
nominal thrust, thereby giving a thrust range of 600 to 6000 N [1].
Detailed Vehicle Descriptions Rescue Rover Page 471
Author: David Wyant
Co-Author: Evan Helmeid
Table 5.8.3-2 Required engine characteristics
Required Achievable
Thrust – Nominal (kN) - - 6.000
Thrust – Min (kN) - - 0.600
Weight – Max (kN) 2.283 - -
Weight – Min (kN) 1.732 - -
T:W Optimal Trajectory >2.0 2.6~3.5
T:W Land <0.7 0.30~1.0
Table 5.8.3-2 summarizes the requirements and the capabilities of the Rescue Rover engine.
We use the requirements to size the engine while considering the impact of the thrust levels on
the propellant mass; the engine meets or exceeds all requirements of thrust levels and ranges.
Using Rocket Propulsion Analysis [2], a software package designed for thermodynamic
analysis, the dimensions and performance parameters of a rocket engine can be found easily.
The target goal for this particular main engine was an Isp of 480 s with a thrust of 6 kN. The
throttled and nominal performance characteristics are seen below.
Table 5.8.3-3 Main engine performance characteristics
Full Throttle 10% Throttle
Thrust 6 kN 0.6 kN
Isp 479.4 s 451.22 s
O/F
C*
Pc
5.731
2318.33 m/s
5 MPa
5.731
2318.33 m/s
5 MPa
Expansion Ratio 233.52 233.52
Detailed Vehicle Descriptions Rescue Rover Page 472
Author: David Wyant
Co-Author: Evan Helmeid
Figure 5.8.3-1 The nozzle geometry of the main engine is shown above. The dimensions are
in mm.
The total mass of the main engine is 12.15 kg (including associated turbo-machinery). It was
found using the process outlined in Space Propulsion Analysis and Design [3].
The main thruster will make use of a liquid hydrogen and liquid oxygen as the fuel and
oxidizer, respectively. These propellants were chosen for their ability to be synthesized on the
surface of Ceres by the ISPP facilities. The total mass of propellant that will be burned over a
flight of the Rescue Rover varies depending upon the mission; however the maximum range
mission requires 2154.7 kg of propellant. This will be expelled over a system of orbit burns, de-
orbit burns and hovering burns totaling 1820 seconds of burn time.
Detailed Vehicle Descriptions Rescue Rover Page 473
Author: David Wyant
Co-Author: Evan Helmeid
5.8.3.2 Attitude and Control Thrusters
The design of our Rescue Rover calls for an attitude and control system to the keep the
Rescue Rover oriented during flight. This system will be comprised of four gimbal-able
thrusters, each capable of delivering 2.2 kN of thrust. The specifics of the system will be
detailed later in the paper, only the thruster characteristics will be addressed in this portion.
Table 5.8.3-4 Thruster performance characteristics
ADCS Thruster
Thrust 2.2 kN
Isp 453.8 s
O/F
C*
Pc
4.98
2369.17 m/s
5 MPa
Expansion Ratio 37.971
Figure 5.8.3-2 The nozzle geometry of the attitude thrusters is shown above. The
dimensions are in mm.
The mass of each thruster is 7 kg (including associated turbo-machinery). This imparts a
total mass of 28 kg to the Rover. The mass was found using the same process as above.
As with the main engine, liquid hydrogen and liquid oxygen will be used as the propellant for
the attitude and control system on the Rescue Rover. This will not only enable the propellant to
Detailed Vehicle Descriptions Rescue Rover Page 474
Author: David Wyant
Co-Author: Evan Helmeid
be produced on the surface of Ceres, but also allows for the sharing of tankage between the
thrusters and the main engine, cutting down on the overall mass of the craft. As with the main
engine, these numbers were found using RPA [2].
Detailed Vehicle Descriptions Rescue Rover Page 475
Author: David Wyant
Co-Author: Evan Helmeid
5.8.3.3 Maneuvering Wheels
For the Rescue Rover to function well as a rescue rover, the ability to dock with the stranded
Exploration Rovers is essential. Therefore, fine maneuvering wheels were added to the Rover.
These wheels allow the Rover to traverse terrain, still able to arrive at the emergency site if it is
unsafe to land there and then position the Rover in an appropriate docking position.
Each landing strut will have a wheel attached to it and each wheel will possess its own electric
motor. The wheels will be identical to those on the exploration rovers. The power requirement
of 4.66 kW allows the rover a nominal operating speed of 2 m/s while driving up a 45 degree
incline. Each motor has a mass of 0.72 kg for a total propulsion system weight of just 2.88 kg
[4]. The top speed was lowered from that of the exploration rovers because the distances the
rover must traverse are much lower. Unlike the other rovers, each motor will be capable of
delivering enough power to move the entire craft, should something happen to the other motors.
The steering for the Rover will be much the same as it is for exploration rovers: each wheel
can turn at an independent speed. Unlike the other rovers, this is a product of the electric motors.
No transmissions are required or necessary on the Rescue Rover.
Detailed Vehicle Descriptions Rescue Rover Page 476
Author: David Wyant
Co-Author: Evan Helmeid
References
[1] "CECE” Pratt & Whitney Rocketdyne, 2009.
[2] Ponomarenko, A. “RPA: Design Tool for Liquid Rocket Engine Analysis.” May 2010.
[3] Humble, R.,Henry, G., Larson, W., Space Propulsion Analysis and Design, Mc-Graw Hill,
1995.
[4] "Launchpoint Motors,” [http://www.launchpnt.com/Documents/Dual%20Halbach-Motor-
Data-Sheet_R1.pdf. Accessed 3/31/11.]
Detailed Vehicle Descriptions Rescue Rover Page 477
Author: Ben Stirgwolt
Co-Author: Jillian Roberts
5.8.4 Human Factors Systems and Habitability Considerations
The Rescue Rover is designed to be used for a day at most. In order to keep the size of the
rover to a minimum, we keep only the absolute necessary items on the rover. There is neither a
lavatory (diapers are used instead) nor a galley (a small amount ready-to-eat food is stored
onboard). There are two medical beds, sink and medical equipment including a defibrillator,
ventilators, and basic First-Aid supplies.
Because the astronauts could spend a day in the Rescue Rover, we must make water
provisions to allow for drinking, food rehydration, and basic hygiene while away from the CTV
for four crew members. We conducted a trade study which determined that storing water,
instead of recycling it, would significantly decrease mass in the rover. This trade study can be
found in the Appendix. The total mass, volume, and power requirements for the Rescue Rover
are found in the table below.
Table 5.8.4-1 Specifications for the water supply and recycling system of the Rescue Rover
Crew
Members
Days Mass, kg Power, kW Volume, m3
Water Supply and
Regeneration
4 2 244.8 0 0.245
Detailed Vehicle Descriptions Rescue Rover Page 478
Author: Ben Stirgwolt
Co-Author: Jillian Roberts
In case of a fire, the Rescue Rover has two fire extinguishers and one smoke detector. See the
Fire Suppression and Detection section from the Crew Transfer Vehicle and its corresponding
Appendix for details. The mass, power, and volume can be found in the table below.
Table 5.8.4-2 Specifications for the fire suppression and detection system
Mass, kg Power, kW Volume, m3
Fire Detection and Suppression 23.27 0.0015 0.0788
To provide an ergonomic working environment which is well-lit, the Rescue Rover will have
a lighting system. The table below describes the mass, power, and volume of the lighting
system. We assume the Rover needs 1000 lux for the astronauts to efficiently perform medical
procedures in case of injury.
Table 5.8.4-3 Specifications for the lighting system
Mass, kg Power, kW Volume, m3
Lighting System 108 0.900 3
Detailed Vehicle Descriptions Rescue Rover Page 479
Author: David Schafer
5.8.5 Attitude Determination and Control Systems
As the Rescue Rover moves using powered flight, it has a specific need to be three-axis
stabilized. In order for us to stabilize the rover, we must implement both an attitude
determination system and a physical actuating system. The determination system has no need
for high accuracy, yet we do need the actuating system to be both very quick and very powerful.
Finally, the rover needs autonomous control capabilities for its transfer from station 1 to station 2
at the midpoint of the mission.
5.8.5.1 Attitude determination
To understand the Rover’s orientation, we need an attitude determination system. This
system consists of two separate inertial reference units, called Motion Reference Units (MRU)
connected to two separate computers. As the determination system is inertial, measuring
changes with accelerometers, the computers are necessary to integrate the information and
produce a mathematical model for how the vehicle has changed orientation. This system is
accurate to 0.02 deg, easily achieving the 0.2 deg needed to communicate with the satellite
network. As with the exploration Rovers, we use a Kongsberg 5+ model MRU as a model for
the sensors, and again, the internal gains must be changed to account for the change in gravity
between Earth and Ceres [1].
Detailed Vehicle Descriptions Rescue Rover Page 480
Author: David Schafer
5.8.5.2 Actuators
We use four stabilizing thrusters mounted on Canfield joints to keep the Rover pointed in the
proper directions. These thrusters are not perfectly positioned for attitude control, as they also
aid in the launches of the craft, giving us some undue translational effects to be covered shortly.
The Canfield joints allow the thrusters full hemispherical, three degrees of freedom to move,
allowing for ideal thrusting to turn the Rover [2].
As the inertias of the craft are relatively large (approximately1.5*104 kg*m2), and the Rover
needs to be stabilized quite quickly, the attitude control thrusters need to be quite large,
producing about 2.2 kN of thrust, each, in any one direction (within their range of directions).
Such high forces allow us to correct the Rover from perturbations in a very short time period.
The large thrusters increase the inert mass of the craft, yet the ability for very short burn times
allows for a low propellant mass for the stabilization maneuvers. Approximately 100 kg of
propellant waits in the tanks, ready to fire the moment the Rover needs it.
We positioned the actuating thrusters only 45° out from the bottom of the craft, instead of 90°,
for two main reasons, as seen below in Fig. 5.8.5.2-1. First, this will allow the thrusters to aid in
the launch of the craft while only putting a low torque on the Canfield joints for the extended
period of the launch, and second, to give the thrusters a bias, or a “zero” position where they are
more able to keep the bottom of the craft pointing down. However, this means that the thrusters
are not a coupled set, but instead will provide translational effects when used to rotate the body.
For instance, during a launching process, the Rover must turn a full 90 degrees. This will cause
the Rover to move at about 0.7 meters per second in the direction directly opposite of where the
Rover is attempting to go. As the magnitude of the final velocity of the launch itself is over 200
meters per second, this translational effect is minimal, and can be neglected once the use of the
Detailed Vehicle Descriptions Rescue Rover Page 481
Author: David Schafer
Rover’s wheels is considered. To correct for the lost velocity, the Rover can either burn
propellant for a fraction of a second longer, or move the forty meters it will have traveled after it
lands, on its wheels.
Figure 5.8.5.2-1 Location of the attitude control thrusters on the Rescue Rover
Another aspect of the Rover’s correctional facilities is how it maneuvers on the ground. As
the gravity is quite low on Ceres, and the Rover will be moving on the surface at a decent
Attitude Thrusters
Detailed Vehicle Descriptions Rescue Rover Page 482
Author: David Schafer
velocity, with a high center of mass (at the end of the mission), the Rover could possibly tip
over. This would be a huge problem, but thanks to the positioning of the thrusters, a torque can
be applied to correct for this without adding any vertical translational effects. The Rover propels
to the side at a low velocity (around 0.3 meters per second, or 0.6 miles per hour), but rights
itself without forcing it neither off of, nor into, the surface of Ceres.
The totals for the entire attitude determination system of the Rescue Rover are presented
below in Table 5.8.5.2-1
Table 5.8.5.2-1 Full mass, power, and volume requirements of the entire attitude
determination and control system of the Rescue Rover
Hardware Mass, kg Power, kW Volume, m3
MRU’s 5 0.024 0.02
Computers 3 0.01 0.004
Propulsion system 28 0
0.002
Propellant 95 0 0.001
Total 131 0.034 0.027
Detailed Vehicle Descriptions Rescue Rover Page 483
Author: David Schafer
5.8.5.3 Autonomous commands
Finally, the autonomous commands must be considered. We move the Rescue Rover from
Station 1 to Station 2 autonomously at the midpoint of the mission, and for this one maneuver a
system must be made to control the Rover. As we already have an inertial system of sensors
combined with a computer, integration of the autonomous controller is a simple aspect of giving
the Rover a course into the controller already designed to control the thrusters. The dual
computer system can control this task with relative ease, but will be looked at further in depth in
another section.
Detailed Vehicle Descriptions Rescue Rover Page 484
Author: David Schafer
References
[1] Kongsberg Maritime “MRU 5+ Datasheet,” Trondheim, Norway
[2] Royer, Caleb, “Robotic Canfield Joint,” National Instruments, April 29, 2010
Detailed Vehicle Descriptions Rescue Rover Page 485
Author: Kim Madden
5.8.6. Structural and Thermal Control Systems
5.8.6.1. Structural Components
Our Rescue Rover consists of a pressure vessel main body, two clear ellipsoidal windshields,
landing legs, propellant tanks and radiation shielding. All of these components come together to
create the general cylindrical shape of the Rover. The structure of the Rescue Rover safely
contains the crew and essential life support systems.
Pressure Vessel
Pressure vessels are traditionally circular, because of the excess stresses introduced by
bending. We choose a cylinder shape in order to get the most space inside. A sphere is the ideal
shape, but needs a very large radius in order to contain the required equipment. With a cylinder,
the radius can be smaller since the length can be changed.
First we need to determine the radius required for the cylinder. Based on the inside
configuration, we need a floor length of 3 m. We decide to place a floor along a chord of the
cross section, instead of across the middle. Having it lower makes the necessary radius smaller,
and gives more head room for a crew member to comfortably stand. Figure 5.8.6.1-1 shows a
circular cross section with the approximate floor location.
Detailed Vehicle Descriptions Rescue Rover Page 486
Author: Kim Madden
Figure 5.8.6.1-1 A circular cross section, with important variables defined. We wish to find
the radius using the desired storage height h and the chord length c.
We choose to have a storage height h of 1 m. This creates enough storage space for the life
support systems and a propellant tank, and minimizes the actual radius of the cylinder. While
calculating the radius, we also ensured that a man would be able to stand up comfortable (called
the head room).
The radius of the cylinder is 1.69 m. With the storage height of 1 m, there is 2.37 m of head
room above the floor, which is enough for a man to stand up. The sides have a smaller height,
but there will be counters and dock doors there, so it will not be an issue. The length of the
cylinder is 3.7 m. This is dictated by the previously designed floor plan.
We make the thickness of the walls to be 1.5 cm. We can now determine the mass of the
pressure vessel part of the Rover by multiplying this calculated thickness by the surface area.
The mass of the pressure vessel is 1,658.25 kg. The internal volume is 33.16 m3.
Windshields
We add windshields to the front and back of the Rovers to serve a number of different
purposes. The first obvious reason is that we want the crew to be able to see where they are
driving. This way, they can avoid rocks or other obstacles, as well as have a good view of the
Ceres surface. Another reason to include them is to keep Ceres dust and dirt out of the Rover,
Diagram by Kim Madden
R
C
h
D
Detailed Vehicle Descriptions Rescue Rover Page 487
Author: Kim Madden
maintaining a clean and safe environment for our crew. Lastly, these end caps for the Rover will
create a closed area to maintain the internal pressure.
We choose to make our windshields out of polycarbonate. It is a stronger and more durable
material than plexiglass. The material needs to be clear so that the crew can see out of it.
Polycarbonate has a yield strength of 62.1 MPa, and a density of 1200 kg/m3 [1].
The shape we choose for the windshields is a 2:1 ellipsoidal head. This shape can be seen in
Fig. 5.8.6.1-2. This design would have the same benefits as the hemispherical shape, but would
reduce extra mass because it does not stick out as far.
Figure 5.8.6.1-2 The 2:1 ellipsoidal windshield, which reduces mass while maintaining the
visibility
We design the ellipsoidal to have a thickness of 1 cm. The surface area of two ellipsoidal
windshields is 15.25 m2. By multiplying the surface area by the required thickness and the
density of polycarbonate, we determine the mass of two ellipsoidal windshields. The mass is
549.36 kg, and the internal volume is 5.02 m3.
Floors
We must include floors so that the crew can stand and work on them. As previously
discussed, we already know the dimensions of the floor, which are based on the configuration.
The floor is a rectangle 1 m above the bottom of the cylinder. In order to determine the thickness
of the floor, we model it as a beam fixed at both ends with a concentrated load in the middle.
2R R/2 R/2
Diagram by Kim Madden
Detailed Vehicle Descriptions Rescue Rover Page 488
Author: Kim Madden
This gives a very conservative estimate of the floor thickness, because all of the mass will not be
concentrated in the middle, but rather, spread around. We assume that 2/3 of the human factors
and science mass is on the floor, while 1/3 of it is in storage. We also assume that the maximum
deflection of the floor can be 1 cm. This deflection occurs in the center [2].
We build the floor to be 2 cm thick. This results in a floor mass of 747.62 kg, and a volume of
0.27 m3.
Radiation Shielding
Because we are traveling in space, we need some radiation shielding to protect the crew and
electronics from harmful exposure. There are many different ideas for the ideal radiation
protection, but it is hard to get an exact value since studies can only be done in space. A heavily
shielded area is needed in case of a solar particle event and galactic cosmic rays. This is located
in the CTV; the Rovers will not require this much shielding.
We choose to use a passive shield consisting of aluminum and polyethylene. Polyethylene
contains a lot of hydrogen and is lightweight, making an excellent shield for radiation. It has a
density of 925 kg/m3
[1]. The outer layer of the Rover is already made of 1.5 cm of aluminum, so
this also doubles as radiation shielding. However, we need a thicker shield to be effective. For a
light shielding, we use 80 kg/m2 of material [3,4]. The aluminum pressure vessel contributes
42.15 kg/m2, so we require an additional 37.85 kg/m
2. To add the additional mass, we need
another 4 cm of polyethylene. This is inside the pressure vessel portion of the Rover, and has a
mass of 1,489.08 kg.
Landing Legs
The Rescue Rover lands on the surface of Ceres with a set of four landing legs. These are
made of carbon fiber. We model the legs with shock absorbers to help provide a soft landing and
Detailed Vehicle Descriptions Rescue Rover Page 489
Author: Kim Madden
to avoid excess stresses on the vehicle. The shock absorber consists of a compressible spring
with the capability of absorbing the shock of a landing at 10 m/s. It is made of a strengthened
steel spring which compresses to absorb landing energy from a hop. The legs are designed by
Andrew Curtiss, and the analysis can be found in Appendix 5.2.7.2.
The mass of four landing legs for the Rescue Rover is 72.90 kg, and has a volume of 0.03 m3.
Propellant Tanks
There are three tanks on the Rescue Rover to store the necessary propellant. The propellant
requirements are discussed in Sections 5.8.3 and 5.8.8.
To store the required 337.4 kg of liquid hydrogen, we need a cylindrical tank with an inside
volume of 4.79 m3. We design this cylindrical tank to fit in the storage area in the Rescue Rover.
The height is 0.16 m, and the radius is 1.35 m. Based on tank sizing designs by Alex Kreul
(Section 5.4.5.4), the thickness of the tanks should be 2 cm. The mass of the LH2 tank is 285.67
kg, and a volume of 1.18 m3. Figure 5.8.6.1-3 shows a diagram of the cylindrical LH2 tank.
Figure 5.8.6.1-3 This is a diagram of the LH2 cylindrical tank, located inside the Rescue
Rover
We also need two hemispherical tanks for liquid oxygen. We place a hemisphere on each side
of the main engine on the bottom of the Rover. In order to store the required 1,933.31 kg of LOX
in each tank, the internal volume must be 0.85 m3. Since we already know the volume, we can
0.16 m
1.35 m
Diagram by Kim Madden
Detailed Vehicle Descriptions Rescue Rover Page 490
Author: Kim Madden
easily calculate the required radius for the hemisphere. The radius is 0.64 m. Figure 5.8.6.1-4
shows a diagram of the LOX tanks.
Figure 5.8.6.1-4 This is a diagram of the location and sizes of the LOX tanks. There is one
on each side of the main engine. They are on the bottom side of the Rover.
The mass of the two hemispherical tanks is 285.67 kg, and the total volume is 1.18 m3.
Adding these together, we know that the total mass of the propellant tanks is 935.35 kg, and the
total volume is 2.32 m3.
Buckling
Now that we have the basic structure of the Rover complete, we need to make sure that it
doesn’t buckle during the launch. We assume that during the STV launch, it will experience 6g’s
of acceleration. The force during launch is the total mass of the Rover times the acceleration
during launch, which is 337.4 kN.
The force that would buckle the Rover during launch is 2,934,132 kN. Since this force is
MUCH larger than the force the Rover will experience during launch, we conclude that the
Rover will not buckle during the launch.
Nuts, Bolts and Screws
In order to account for various building materials, such as nuts, bolts and screws, we add 10%
of the total structure mass to the totals, as well as 5% of the structural volume. These are
Bottom of Rescue Rover 0.64 m
0.64 m
0.64 m
0.64 m
LOX Tank LOX Tank
Main Engine
Diagram by Kim Madden
Detailed Vehicle Descriptions Rescue Rover Page 491
Author: Kim Madden
approximated values, and while they may seem negligible, they actually add up to 444.43 kg and
0.15 m3.
Structural Summary
Table 5.8.6.1-1 shows a summary of the mass, power, and volume requirements of the
structural components of the Rescue Rover. Figure 5.8.6.1-5 shows a picture of the Rescue
Rover, with the structural components pointed out.
Table 5.8.6.1-1 Structural summary of mass, power and volume parameters
Component Mass, kg Power, kW Volume, m3
Pressure Vessel 1,658.25 0 33.16
Windshields 549.36 0 5.02
Floors 747.61 0 0.27
Radiation Shielding 1,489.08 0 1.61
Landing Legs 72.90 0 0.03
Propellant Tanks 935.35 0 1.14
Nuts, Bolts and Screws 444.43 0 0.15
Totals 4,961.64 0 38.90
Detailed Vehicle Descriptions Rescue Rover Page 492
Author: Kim Madden
Figure 5.8.6.1-5 Depiction of the Rescue Rover, with the structural components pointed
out. All other components are located on the inside of the Rover
Pressure Vessel
Windshields
Landing Legs
LOX Tanks
Detailed Vehicle Descriptions Rescue Rover Page 493
Author: Kim Madden
References
[1] Callister, W. D., Materials Science and Engineering An Introduction, 7th
ed., John Wiley &
Sons, Inc., Pennsylvania, 2007, Appendix B.
[2] Gere, J. M. and Goodno, B. J., Mechanics of Materials, 7th
ed., Cengage Learning, Ontario,
2009, Chaps. 5, 6, 8, 11.
[3] Wilson, J. W., Miller, J., Konradi, A., and Cucinotta, A. F., “Shielding Strategies for
Human Space Exploration”, National Aeronautics and Space Exploration, December, 1997.
[4] National Council on Radiation Protection and Measurements, NCRP Report No. 98:
“Guidance On Radiation Received In Space Activities”, Bethesda, MD: NCRP, 1989.
Detailed Vehicle Descriptions Rescue Rover Page 494
Author: Kim Madden
5.8.6.2. Thermal Control Systems
A thermal control system is important for all spec vehicles, especially manned vehicles. While
space is quite cold, there are electronics and motors inside each vehicle that produce heat. There
needs to be a way for the heat to escape the vehicle so it doesn’t get too hot.
There are two main sources of power that add heat to the inside of the vehicle. The first is the
heat that is produced from the crew inside. A person produces 61.3 watts of heat. Multiply this
number by the number of people inside the vehicle at any time to get the amount of heat that
needs to be rejected. For the Rescue Rover, there are a maximum of 6 crew members inside at
any time, so 367.8 watts of heat needs to be rejected. Power also comes from the rejected heat
from the electronics, which produced because the electronic components are not 100% efficient.
We assume that the electronics were 65% efficient, as advised by Dr. Boris Yendler. The
electronics in the Rescue Rover require 2.414 kW of power, so 0.973 kW need to be rejected.
There are two ways that heat leaves the Rescue Rover. The first is due to the colder
temperatures on Ceres. During a Ceres day, the temperature on the surface is 235 K, and it is 100
K during the night. We want to maintain the inside of the Rover at a comfortable temperature for
the crew. We choose to keep it at 293 K, which is a comfortable room temperature on Earth.
Heat will escape the Rover because of the difference in temperatures. We require an additional
heat rejection system, such as radiators and heat pumps. Heat pipes carry heat from the
electronics to radiators on top of the Rover, which then reject excess heat. Figure 5.8.6.2-1 shows
a schematic of the thermal control systems.
Detailed Vehicle Descriptions Rescue Rover Page 495
Author: Kim Madden
Figure 5.8.6.2-1 Schematic of the thermal control systems. The top portion shows what
kind of heat transfer goes into and out of the vehicle, and the bottom portion shows the
inside system.
Multilayer Insulation
We want to minimize the amount of heat lost through the Rover due to environmental
differences. We wrap the Rover in multilayer insulation (MLI) in order to stop some of this heat
flow. MLI blankets are 30 layers of 0.25 mm thick metalized Mylar sheets separated by a mesh.
This acts as a barrier for the heat radiated from the surface of the spacecraft into the cold space.
The outer layer is thicker since it will be exposed to the elements, and white to reflect sunlight
[1]. In order to determine the mass of the MLI covering the Rover, we multiply the surface area
of the Rover that will be covered in MLI (39.342 m2) by the density of MLI.
Heat Pipes
Heat pipes will run all through the Rescue Rover in order to carry heat from the electronics
to the radiators. We choose to use water as the working fluid for the heat pipes. Ammonia is the
traditional working fluid; however, if there was a leak, the crew would be in trouble. Water will
Diagram by Kim Madden
Detailed Vehicle Descriptions Rescue Rover Page 496
Author: Kim Madden
be available on the surface, so if it needs to be replenished, it can be easily. Also, a water leak
will not harm the crew in any way.
When the water flows under an electronic, it will heat up and vaporize. As the water moves
away from the electronics towards the radiators, the water will condense. This is how heat moves
throughout the pipes. Small resistance heaters are located near the radiators to keep the water
liquid when it starts to get colder. If the water freezes when exposed to the radiators, the heat
pipe would then be useless and the Rover will overheat. The mass of the heat pipe, including the
water required, is 25.301 kg.
Radiators
We must now determine the mass and size of the radiators. The radiators are located on top
of the Rovers. The radiators will be required to open and close depending on how much heat
needs to be rejected. Rubber corners connect the heat pipes through the radiators to the heat
pipes in the vehicle. This allows the radiators to fold, and also stops the flow of water when the
radiators are folded. They can be closed during the night to stop heat flow to keep the inside
warm, and then open up again during the day. On the Rescue Rover, there will be 4 sets of 2
radiators (8 total radiator panels). One side of each radiator set will be covered in MLI to stop
more heat flow. This leaves ¾ sides of each radiator set to radiate heat. The power requirements
for this mechanism can be found in section G.3.2 by Joel Lau. The radiators require 0.016 kW of
power to fold open and close.
To remove a certain amount of heat, the radiator needs to have a certain surface area. The size
of one radiator panel is 0.293 m by 0.293 m, giving a surface area of 0.6844 m2. The mass of all
the radiator panels is 184.857 kg. This includes the MLI covering ¼ of the surface area of a
radiator set.
Detailed Vehicle Descriptions Rescue Rover Page 497
Author: Kim Madden
Aluminum Plates
For heat to be transferred to the heat pipe from the electronics, an aluminum plate needs to be
underneath. We assume that there is 0.5 square meter of aluminum throughout the Rover. This is
broken up and placed under every electronic, with the heat pipes flowing under the plate. The
thickness of the plate is 5 mm.
Heater
We have decided not to include a heater in the Rescue Rover. Ideally, the Rover will never
need to be used; but when it is, it will only be for short amounts of time (under 4 hours). While
the Rescue Rover is docked to the CTV, awaiting a Rescue mission, it will have the same
equilibrium temperature as the CTV, 293 K. There are still electronics running inside, as well as
up to 6 crew members, adding heat to the inside of the Rover.
Thermal Control Summary
Table 5.8.6.2-1 shows a compiled chart of the mass, power, and volume requirements for the
Rescue Rover thermal control system.
Table 5.8.6.2-1 Rescue Rover thermal control system summary
Component Mass, kg Power, kW Volume, m3
MLI Covering 10.976 0 0.039
Heat Pipe 25.301 0 0.218
Radiators 184.857 0.016 0.066
Aluminum Plates 7.025 0 0.003
TOTAL 228.160 0.016 0.325
Detailed Vehicle Descriptions Rescue Rover Page 498
Author: Kim Madden
References
[1] Birur, G. C., Siebes, G, and Swanson, T. D., “Spacecraft Thermal Control”, Encyclopedia
of Physical Science and Techonology, 3rd
ed., Academic Press, 30 March 2001.
Detailed Vehicle Descriptions Rescue Rover Page 499
Author: Tony D’Mello
5.8.7 Communication Systems
5.8.7.1 Internal Communication
In Table 5.8.7.1-1, we show the communication devices that we provide in the Rescue Rover.
Like our Exploration Rovers, our Rescue Rover contains only two television monitors. The
rover has a small antenna located on the ceiling in the center of the rover that can accommodate
communication among six cell phone devices.
Table 5.8.7.1-1 Internal communication device characteristics for the Rescue Rover
2 Televisions 6 Cell Phones 1 Antenna
Mass, kg 10 3.6 1.7e-3
Power, kW 0.2 0.42 1.5e-3
Volume, m3
0.016 0.003 1.94e-7
Detailed Vehicle Descriptions Rescue Rover Page 500
Author: Tony D’Mello
5.8.7.2 External Communication
Because the Rescue Rover must send and receive the largest number of high definition
television (HDTV) channels during times of crisis, we design the Rescue Rover to perform at the
highest frequencies among the radio frequency (RF) connections. We can designate six of the
seven channels for crew communication while the other channel is for the camera that will focus
on the rover in need of rescue. Table 5.8.7.2-1 below displays the mass, power, and volume
values for the transmitter and the receiver.
Table 5.8.7.2-1 External communication device characteristics for the Rescue Rover
1 Transmitter 1 Receiver
Frequency, GHz 60 17
Data Rate, HDTV channels 7 6
Mass, kg 0.37 5.92
Power, kW 0.1 __
Dish Diameter, m 0.7 1.5
Detailed Vehicle Descriptions Rescue Rover Page 501
Author: Evan Helmeid, Megan Sanders
5.8.8 Trajectory and Flight Path
One of the most important factors we consider to determine the trajectory of the rescue rover
is the uncertainty in the surface features of Ceres, especially when traveling the maximum
distance. Due to current image resolution of the destination, we may encounter surface features
up to 18 kilometers in altitude. As such, one of our objectives is to reach and maintain a height of
more than 18 kilometers before traveling too far downrange from the launch site (see appendix
F.4.1.1 for an explanation of the 18-km restriction).
As a result of the surface feature uncertainty the rescue rover uses an optimal launch and
landing trajectory to a transfer orbit at 25 km altitude. The rover never completes a full orbit
about Ceres. Instead the rover de-orbits to land at the location of the stranded astronauts and
picks them up. The rover then launches and transfers the astronauts back to the Crew Transfer
Vehicle (CTV) location.
This solution is not feasible for all scenarios because the optimal trajectory requires a
minimum downrange distance to achieve orbit, rotate 180° and then land at the target site. If the
target is closer than 207.5 km, then the rover uses a suborbital ballistic trajectory to prevent
landing beyond the target. If the target is within a radius of 7.2 km from the CTV location, the
rover drives to the target location and back. The rover can actually drive a maximum distance of
9.3 km over flat ground but we restrict the distance to 7.2 to ensure that if any surface
irregularities are encountered the rover will be able to still reach its destination.
Trip distance, times, and propellant requirements are summarized in Table 5.8.8-1.
Detailed Vehicle Descriptions Rescue Rover Page 502
Author: Evan Helmeid, Megan Sanders
Table 5.8.8-1 Rescue rover trip specifications depending upon distance to travel
Method
Optimal to
transfer orbit
Ballistic
trajectory Drive
Units
Downrange distance 207.5~765.4 7.2~207.5 0.0~7.2 km
Propellant mass 2155 319~2025.2 0 kg
Total trip time 31.33~88.87 11.13~60.53 0.0~60 minutes
In our propellant mass and total trip time we account for a 60-second hover on both ends of
the trip, once for landing at the target location and once for landing at the CTV location after the
rescue. This hover is built-in to allow for inaccuracies in the control system and to find an
appropriate landing location for the astronauts.
While running several case options, we vary the Isp and thrust inputs. These changes affect
the total amount of propellant necessary, and we select the trajectory that uses the least amount
of propellant. With the chosen trajectory, we still require small tanks on the outside of the
vehicle but we are able to keep these external tanks to a minimum.
Detailed Vehicle Descriptions Rescue Rover Page 503
Author: David Schafer
5.8.9 Autonomous Operations
In order for the crew to transfer from Station 1 to Station 2 at the midpoint of the mission, we
need the Rescue Rover to cross over Ceres without any manned guidance. This means an
autonomous system must guide the Rover halfway across Ceres. This system must be fully
integrated into the positioning, power, propulsion, and even the communications systems in
order to operate the Rover on its solo journey.
The Rescue Rover will travel to ISPP Station 2 by launching itself into orbit, completing half
an orbital revolution and then de-orbiting to land at the new Station. This transfer takes a similar
amount of time as the CTV transfer, but the Rescue Rover arrives at the new site at a slightly
later time, making it available to go rescue one of the exploration Rovers if necessary.
As the Rescue Rover launches and lands on Ceres itself, spending the rest of the time above
the surface of Ceres, only small changes must be made to the current systems to prepare it for
autonomous control. We have already set up the attitude control system to automatically keep
the vehicle stable throughout its missions (see appendix section A5.8.5.5), as well as the
communications system to stay in contact with the satellite network. Additionally, the power
systems are already integrated into the computer network, so the system need only add the
launching and landing considerations.
Launching is the easy part of the mission, where simply the coordinates of the first landing
zone must be used to find the specific trajectory the Rover will take. This information is then
input to the thruster system which takes the Rover exactly where it needs to go. Then, once the
Rover has reached the first landing zone, the Rover must simply repeat this process to launch
towards the second station. Landing at each of these zones is a bit trickier. Here we will have
set up a specific coordinate for the Rover to land at, and the Rover will have to use the already
Detailed Vehicle Descriptions Rescue Rover Page 504
Author: David Schafer
formulated landing system to slow down the descent and to ensure stability whilst landing. This
method implies both that the two landing zones will already have been investigated, and second,
that the astronauts will have already made it to the second Station before the second launch of
the Rescue Rover. In other words, we require specific information about the landing zones of the
Rover in order to use this method of landing the Rover.
Other considerations for the autonomous controls of the Rescue Rover are presented in
Appendix C.
Detailed Vehicle Descriptions Rescue Rover Page 505
Author: Megan Sanders
5.8.10 End of Life Configuration
We choose to end the useful life of the Rescue Rover by crashing it into the surface of Ceres.
This crash provides the necessary force to complete the calibration of the seismic stations. The
Rescue Rover is filled with propellant before using its autonomous system to launch vertically
into the air. The engine burns for 21 seconds before shutting off and the vehicle continues to
coast upwards. The Rover reaches a maximum height of 26.58 km above the surface before the
pull of Ceres’s gravity becomes too great and the Rover is pulled back down. Just prior to the
maximum height the attitude control motors are used to rotate the Rover so that the main engine
is oriented facing away from the Ceres surface. Once the Rover reaches its maximum height, the
main motor will be fired, accelerating the Rover towards the surface of Ceres. The engine
continues to fire until the Rover contacts the surface in order to produce the maximum amount of
acceleration. The pertinent values at the time of impact are given in Table 5.8.10-1.
Table 5.8.10-1 Rover specifications at moment of impact
Parameter Value Unit
Total Mass 8,339 kg
Unused Propellant 344.4 kg
Acceleration 5.99 m/s2
Force 49,962 N
We opt to burn the engines for only 21 seconds, rather than completely use all of the
propellant for two main reasons. The first reason is that, by burning for a shorter time, the
Rover’s maximum height will be reduced. Keeping the maximum height low will allow the
gravity of Ceres to continue to act on the Rover and pull it down as expected. The second reason
is that, by not using all the possible propellant in the initial launch, there will be unused
propellant left to use for a second burn. The second burn allows us to accelerate towards the
Detailed Vehicle Descriptions Rescue Rover Page 506
Author: Megan Sanders
surface of Ceres. There is still propellant remaining in the Rover when it impacts the surface of
Ceres. The propellant left provides an increase in mass that allows for a greater impact force into
Ceres.
A more detailed discussion of the thought process behind this configuration and the algorithm
used to calculate it is located in Section A.5.8.10.
Detailed Vehicle Descriptions Communications Network Page 507
Author: Justin Axsom
5.9 Communications Network
5.9.1 Ceres Orbiting Satellites (COS)
In order to relay data between Ceres and Earth, we employ the use of the Ceres Orbiting
Satellites. The COS handle all of the telecommunications and logistics data for Ceres based
operation through the use of optical and radio frequency communications technology.
Detailed Vehicle Descriptions Communications Network Page 508
Author: Graham Johnson
Co-Author: Sarah Jo DeFini
5.9.1.1 Trajectory
Nomenclature
L1/L2 = Libration Points
ΔV = Delta-V imparted, km/s
mpropellant = Mass of Propellant required, Kg
LCO = Low Ceres Orbit, 50 km
ECCO1(2) = Earth Ceres Communication Orbiter
Model Assumptions
The manifold transfer trajectory chosen was located at an altitude of 50.9 km. We
assume that ECCO 1 and 2 are in LCO altitude exactly matching the desired transfer
trajectories.
The ΔV maneuvers are assumed to be impulsive, thus we neglect burn times and burn arc
calculations.
Trajectory Overview
Once STV2 captures into Low Ceres Orbit (LCO), ECCO 1 and 2 jettison from the payload
bay and stay in a circular orbit about Ceres. We then place the halo orbiting satellites, ECCO 1
and 2, in their respective halo orbits about the Sun-Ceres Libration points L1 and L2
respectively. This transfer is completed by performing a ΔV maneuver from LCO to each of the
transfer manifolds. See Table 5.9.1.1-1 for magnitudes of the impulses.
Detailed Vehicle Descriptions Communications Network Page 509
Author: Graham Johnson
Co-Author: Sarah Jo DeFini
Table 5.9.1.1-1 ΔV maneuver costs for each halo orbit transfer
Manifold to L1 Manifold to L2
ΔV, Km/s
mpropellant, Kg
0.604
2897.5
0.2508
1137.8
After each of the ΔV maneuvers, ECCO 1 and 2 transfer on their respective manifold surfaces
for 681.3 and 677.7 days respectively. Each of the transfer manifolds is stable and
asymptotically approaches the desired halo orbit. This allows for a zero ΔV cost to insert each
satellite into their respective halo orbits. A figure of all possible transfer manifold surfaces is
shown in Fig. 5.9.1.1-1. The blue manifold surface represents all the possible transfer
trajectories to L1, the red to L2, and the blue circle around Ceres represents its sphere of
influence.
Detailed Vehicle Descriptions Communications Network Page 510
Author: Graham Johnson
Co-Author: Sarah Jo DeFini
*Courtesy of Christopher Spreen; Purdue University
Figure 5.9.1.1-1 All possible manifold transfer orbits to L1 and L2
The actual transfer manifolds we choose, keep the ΔV s nearly tangential at the manifold
intersection with LCO. This choice helps to keep the ΔV s low in cost. Two scaled figures of the
final halo orbits are shown in Figs. 5.9.1.1-2 and 5.9.1.1-3.
Once the ECCO 1 and 2 reach their halo orbits, we leave them to orbit L1 and L2 for the
duration of the mission, each with an orbital period of approximately 832 days. For the
remainder of the mission, each satellite will help to create a communication link between all
Ceres operations and Earth.
0.9995
1
1.0005
-2
0
2
x 10-4
-1
0
1
2
x 10-4
Nondimensional X-Direction
Ceres Halo Orbit Transfer Manifolds
Nondimensional Y-Direction
Nondim
ensio
nal Z
-Direction
Detailed Vehicle Descriptions Communications Network Page 511
Author: Graham Johnson
Co-Author: Sarah Jo DeFini
*Courtesy of Michael Mueterthies; Purdue University.
Figure 5.9.1.1-2 Overall view of scaled halo orbits for ECCO 1 & 2 with Sun in –x direction
-500
0
500
-200
0
200
-100
0
100 L2
x (Ceres Radii)
Ceres
Final Halo orbits for ECCO 1 & 2
L1
y (Ceres Radii)
z (C
eres
Rad
ii)
Detailed Vehicle Descriptions Communications Network Page 512
Author: Graham Johnson
Co-Author: Sarah Jo DeFini
*Courtesy of Michael Mueterthies; Purdue University
. Figure 5.9.1.1-3 Side view of scaled halo orbits for ECCO 1 & 2 with the Sun left of L1
We employ occasional station keeping maneuvers to prevent the orbits from being perturbed
due to gravitational attractions from nearby bodies. Both satellites have added propellant
budgets to occasionally fire the engines and correct for any deviations from the nominal halo
orbit. A table of the largest possible gravitational perturbation forces the satellites are subjected
to, which occurs when Jupiter is in direct alignment with Ceres, is shown in Table 5.9.1.1-2.
These costs are calculated for a five year duration.
Table 5.9.1.1-2 Perturbation force corrections on L1 & L2 due to Jupiter interaction
Jupiter on L2 Jupiter on L1
ΔV, Km/s
mpropellant, Kg
0.2995
1248.3
0.2991
1246.6
Detailed Vehicle Descriptions Communications Network Page 513
Author: Graham Johnson
Co-Author: Sarah Jo DeFini
While other gravitational influences were investigated, a Jupiter-Ceres alignment produced
the largest possible station keeping costs. Therefore it is the only gravitational influence
necessary to account for. To implement the corrective maneuvers we split up the ΔVs into small
impulses which are occasionally fired throughout the mission.
Other station keeping costs are associated with attitude control and specific sensor pointing
and are explained in the Attitude Determination and Control Systems Section 5.9.4.1. For
trajectory designs previously considered, background information of halo orbits, and
justifications please see the appendix Section A.5.9.1.1.
Detailed Vehicle Descriptions Communications Network Page 514
Author: Elle Stephen
5.9.1.2. Power Systems
The Ceres Orbiting Satellites employ solar energy to generate the power necessary for
operation. To harness the solar power, we chose a ZTJ photovoltaic cell with 42.3% efficiency
[1, 2]. These cells have a mass to area ratio of 0.84 kilograms per sq. meter [2].
Because the satellites operate substantially further away from the Sun than Earth orbiting
satellites, the solar irradiation is roughly 1/9 of that found near Earth of 1353W/m2. Taking into
account the decrease in solar irradiation and the photovoltaic cell efficiency, the solar arrays can
only output 63.5W/m2.
Given the minimum power requirement of 53kW, the arrays output 53.7kW of power with an
area of 844.8m2 and mass of 709.6 kg. We create this area by combining sixteen isosceles
triangular sections into a circular formation. The isosceles triangles have a height of 16.5m and
length of 6.4m. When fully deployed, the solar array has a diameter of 33m.
For the duration of the mission, the solar array should be in constant sight of sunlight.
However, if something should happen and the array is blocked from direct sunlight, a Li-ion
rechargeable battery compensates by providing the necessary power.
Detailed Vehicle Descriptions Communications Network Page 515
Author: Elle Stephen
References
[1] “ZTJ Photovoltaic Cell,” Datasheet: Space Photovoltaics, Emcore: Empower With
Light 1984.
[2] "Insolation and Total Solar Irradiance." World of Earth Science. Ed. K. Lee Lerner and
Brenda Wilmoth Lerner. Gale Cengage, 2003. eNotes.com. 2006. 2 Apr, 2011 URL:
http://www.enotes.com/earth-science/ insolation-total-solar-irradiance [cited 18
January 2011].
[3] “Spire Announces World’s Most Efficient Concentrated PV Solar Cell,” Energy Boom,
6 October 2010, http://www.solarfeeds.com/energy-boom/14566-spire-announces-worlds-most-
efficient-concentrated-pv-solar-cell [cited 18 January 2011].
Detailed Vehicle Descriptions Communications Network Page 516
Author: Kyle Svejcar
Co-Author: Graham Johnson
5.9.1.3. Propulsion Systems
We place the Ceres halo orbiting satellites into their respective orbits with a single burn,
having a burn time of 213s each. The satellites (ECCO 1 and 2) take 1.87 and 1.86 years
respectively to transfer into the desired halo orbits.
For both of the Ceres halo orbiting satellites, we design an engine to generate a total thrust of
1447 N with a corresponding Isp of 330s at vacuum conditions. Our system expends a
bipropellant mixture with Monomethylhydrazine (MMH) as the fuel and Nitrogen Tetroxide
(N2O4) as the oxidizer at an oxidizer to fuel mixture ratio of 2.18:1. Our nozzle has an expansion
ratio of 58.12.
Our nozzle keeps cool by radiation cooling. Radiation cooling works when the nozzle heats
up during combustion structure gets red or white hot and heat radiates into space, keeping the
chamber at a reasonable temperature [1].
The fuel feeds from the tanks to the nozzle using a pressure fed system due to the tanks being
less than ten m3. A helium gas supply pressurizes the propellant tanks forcing fuel and oxidizer
to the combustion chamber. Because this combination is considered to be hypergolic, the
propellant mixture will combust on contact, eliminating the requirement of an igniter.
5.9.1.3.1 Inert Mass and Volume Breakdown
The fuel, oxidizer, and pressurant tanks are made out of carbon fiber composite. Tables
5.9.1.3.1-1 and 5.9.1.3.1-2 show the breakdown of our masses and volumes for ECCO 1 and
ECCO 2 respectively.
Detailed Vehicle Descriptions Communications Network Page 517
Author: Kyle Svejcar
Co-Author: Graham Johnson
Table 5.9.1.3.1-1 Mass and volume breakdown for ECCO 1
Component Mass, kg Volume, m3
MMH 756.7 --
N2O4 1650.6 --
Helium Pressurant 40.4 --
Nozzle 10.3 7.7
MMH tank 5.4 0.86
N2O4 tank 6.6 1.4
Pressurant tank 5.9 1.1
Feed system 17.0 --
Support Structure 2.8 --
Table 5.9.1.3.1-1 Mass and volume breakdown for ECCO 2
Component Mass, kg Volume, m3
MMH 297.1 --
N2O4 648.1 --
Helium Pressurant 40.4 --
Nozzle 10.3 7.7
MMH tank 3.9 0.33
N2O4 tank 4.7 0.56
Pressurant tank 5.9 1.1
Feed system 17.0 --
Support Structure 2.4 --
5.9.1.3.2 Nozzle Dimensions
Figure 5.9.1.3.2-1 shows the general shape of our nozzle, while Table 5.9.1.3.2-1 gives the
actual dimensions for the Ceres halo orbiting satellite’s nozzle.
Figure 5.9.1.3.2-1 General shape of nozzle of satellite motor
By: Kyle Svejcar based on
drawing in Rocket
Propulsion Analysis [2].
Detailed Vehicle Descriptions Communications Network Page 518
Author: Kyle Svejcar
Co-Author: Graham Johnson
Table 5.9.1.3.2-1 Dimensions and parameters of Ceres Orbiting Halo Satellits
Variable Length, mm Angle, deg
Rc 30.8 --
R2 80.6 --
Lc 90.3 --
Rt 7.7 --
Rn 2.9 --
Le 152.9 --
Re 58.8 --
Rl 11.5 --
b -- 30
Tn -- 33.6
Te -- 10.3
5.9.1.3.3 Conclusion
The total mass of our propulsion system for ECCO 1 and ECCO 2 are 2945.6 kg and 1182.8
kg respectively and the volumes are 11.0 m3 and 9.7 m
3.
Detailed Vehicle Descriptions Communications Network Page 519
Author: Kyle Svejcar
Co-Author: Graham Johnson
References
[1] Humble, Ronald W., Gary N. Nelson, and Wiley J. Larson. Space Propulsion Analysis and
Design. 1st Ed., Revised. McGraw-Hill, 1995.
[2] Ponomarenko, Alexander, “RPA: Tool for Liquid Propellant Rocket Engine Analysis,” 2010.
Detailed Vehicle Descriptions Communications Network Page 520
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
5.9.1.4. Attitude Determination and Control
The Halo Orbiting Satellites, which provide the communication link between Ceres
operations and Earth, need to have a high degree of stability and accuracy in their pointing.
While the positioning is already highly stable due to their halo orbits, the other environmental
forces can perturb the satellites off of these orbits. These satellites must point towards Earth
with an accuracy of about 400 micro radians (less than 0.02°), as well as simultaneously point at
Ceres and at an oncoming crew (when the crew is en route to Ceres) to effectively create a strong
communication link. Finally, we require the satellites to do all this with a limited amount of
power generated by each satellite’s solar array. As it is so far from the sun, solar power – our
only real option for continuous power of this magnitude – becomes very expensive mass and
volume-wise, and is further explained in Section 5.9.1.3.
5.9.1.4.1 Attitude Positioning
After we place these satellites onto their Halo orbit manifolds, the positioning problem for the
satellites simply becomes knowing where the satellites are, and keeping them there. We attain
the knowledge on the specific locations of the satellites through the use of Motion Reference
Units (MRU’s), an inertial referencing system that uses actuators to find the forces acting on the
body. We require a single MRU triplet to find accelerations in all directions. We must integrate
this information to find velocities and the ultimate position of the satellites, so an onboard
computer must be coupled with this system in order to find exact values of position and velocity
at any given point in time and to map out where exactly the satellite is. The requirement of
tracking all the other vehicles and structures already calls for a computer, so implementing this
MRU system adds no extra mass other than the MRU itself.
Detailed Vehicle Descriptions Communications Network Page 521
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
An over-estimation of the environmental and gravity-gradient forces present in the halo orbits
of Ceres is given in appendix A.5.9.1.4. Ultimately, this analysis gives the satellites a total force
of about 0.005 N. This force can be applied in any direction, and the satellite must counteract
this force. We use our one thruster on the satellite that initially put it into its orbit to correct for
these forces. As the force of the thruster greatly outweighs the environmental forces on the
satellite, we periodically fix for the deviations the satellite has experienced. These deviations
can easily be tracked using the MRU and computer system. Yet, as we only use one thruster to
correct for the external forces, we need to point the thruster in the proper direction first. We
accomplish thruster positioning during the same allotted time for the saturation correction given
below in Table 5.9.1.4.2, and the method of torqueing the satellite to properly point the thruster
is also given below. As shown in Table 5.9.1.4.2-1, the propellant mass is actually quite large,
and this is estimated by the amount the thrusters can be scaled down, as well as the consideration
to operate the satellite for 5 years.
For both of the Ceres halo orbiting satellites, we design an engine to generate a total thrust of
0.1 N but is throttled down to the correct thrust needed, measured by the attitude determination
system, with a corresponding Isp of 200s at vacuum conditions. Our system expends a
monopropellant with hydrogen peroxide (H2O2) as the propellant. Our nozzle has an expansion
ratio of 15.1. The attitude control nozzle keeps cool by radiation cooling in the same manner as
the orbit transfer nozzle. Fuel is fed by the means of a pressure fed system but no pressurant
tank is required for this system.
Detailed Vehicle Descriptions Communications Network Page 522
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
Inert Mass and Volume Breakdown
The propellant tank is also made out of carbon fiber composite for this system. Table
5.9.1.4.1-1 shows the breakdown of our masses and volumes for the attitude control for Halo
Orbiting Satellites.
Table 5.9.1.4.1-1 Mass and volume breakdown for attitude control for the Halo Orbiting
Satellites
Component Mass, kg Volume, m3
H2O2 1248 --
H2O2 tank 5.1 0.73
Feed system 4.4 --
Nozzle 0.69 0.06
Support Structure 0.58 --
Nozzle Dimensions
Table 5.9.1.4.1-2 gives the dimensions for the Ceres Halo Orbiting Satellite’s nozzle based on
dimensions in Fig. 5.9.1.3.2-1.
Table 5.9.1.4.1-2 Dimensions for attitude system
Variable Length, mm Angle, deg
Rc 1.95 --
R2 5.03 --
Lc 64.28 --
Rt 0.48 --
Rn 0.18 --
Le 4.51 --
Re 1.985 --
Rl 0.72 --
b -- 30.0
Tn -- 23.48
Te -- 10.44
The total mass of our attitude system for each Halo Orbiting Satellite is 1258.9 kg and the
volume is 0.79 m3.
Detailed Vehicle Descriptions Communications Network Page 523
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
5.9.1.4.2 Attitude Pointing
We can split our satellite pointing control issues into two main problems: sensing direction
with great accuracy, and changing pointing direction while maintaining this accuracy. Each
aspect introduces its own problems, yet they are all remotely easily solvable.
Attitude determination
We outfit the Halo satellites with the Fine Guidance Sensors currently in use on the Hubble
telescope. Three of these can determine the current position of our satellite, and they can also
find the direction in which each aspect of the satellite points. Finally, they can do this to an
accuracy of about 0.05 micro radians – far within our budget [1]. The only downside of these
sensors is their power requirement. They need about 20 watts each, making a total of 0.06
kilowatts for this part of the system. We need to add a computer to this system in order to map
out and combine all the information from these sensors, as well as track each of the other
satellites, the transfer vehicles, the rovers, the harvesters, and the stations. This bumps the mass
and power requirements only marginally. Combined, this system can map the directions the
satellites are pointing in as well as all the various other vehicles and structures throughout the
mission.
Actuation
Of the two continuous actuating systems that can stay within 0.02 micro radians of accuracy,
the Control Moment Gyroscopes (CMG’s) require the least amount of power. According to the
forces given by the gravity-gradient and environment, the torques that our satellites encounter are
very small in magnitude – only about 0.135 Newton-meters – and thus we only need a small
CMG system to counteract the torques. Creating a system to counteract about 1.2 Newton-
Detailed Vehicle Descriptions Communications Network Page 524
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
meters, we produce the values given below in table 5.9.1.4.2-1. Using a system comprised of
only CMG’s brings up issues with saturation.
Table 5.9.1.4.2-1 Full mass, power, and volume requirements of the entire attitude control
system on the Halo Orbiting Satellites
Hardware Mass, kg Power, kW Volume, m3
MRU 2.5 0.012 0.01
Fine Guidance Sensor 660 0.06 1.275
CMG’s 20 0.8 0.0062
Propulsion System 10.9 0 0.8
Propellant 1250 0 0.128
Saturation effects
As the actuating system is comprised of only CMG’s, the gyros will actually all reach a point
where they are only counteracting the movement of the other gyros. At this point, no torque is
available to turn the satellite, resulting in loss of control of the satellite. This point is known as
saturation of the CMGs, and has been widely investigated across the aerospace community. Our
plan implements a simple computer logic (currently at a Technology Readiness Level (TRL) of
around 6) to take a specific gyro and perturb it once every 4.5 hours (when the other halo
satellite is in use due to the length of the Ceres day), causing the other gyros to correct for the
perturbed one as well as the torques on the satellite [2]. Perturbing this gyro enough, we can
bring this set of CMG’s back to its original position, and release it to again correct for the
torques. This entire process will take an approximated 2.7 hours, and will cause the satellite to
misalign itself from its desired state, causing a loss in communications. This loss is
unacceptable, so this process needs to be done only when the satellite is not needed – which is
when the other satellite is in contact between Earth and Ceres and this specific one is out of
phase.
Detailed Vehicle Descriptions Communications Network Page 525
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
Another problem is presented when the astronauts switch from station 1 to station 2, taking a
full day and needing both satellites. For this single Ceres day (9 hours), the system can more
than easily handle the saturation loads, as saturation occurs only after much longer periods of
time.
The final saturation issue occurs when the communications network needs to add in the move
to use the Relay Satellite. This slew maneuver can be done in about 1.3 hours, meaning that in
the 4.5 hours each satellite will have to correct for saturation and slew over to the relay satellite,
0.5 hours will be remaining time to overlap between satellites (when both satellites are pointing
at the crew).
Detailed Vehicle Descriptions Communications Network Page 526
Author: David Schafer
Co-Author: Graham Johnson, Kyle Svejcar
References
[1] “Instrument Capabilities” Fine Guidance Sensor Data Handbook. Space Telescope Science
Institute, Hubble Division, Baltimore Maryland
[2] Bedrossion, Nazareth. Bhatt, Sagar. Alaniz, Abrin. McCantz, Edward. Nguyen, Louis. and
Chamitoff, Greg. “ISS Contingency Attitude Control Recovery Method For Loss Of
Automatic Thruster Control” American Astronautics Society. February 1-6, 2008
Detailed Vehicle Descriptions Communications Network Page 527
Author: Leonard Jackson
5.9.1.5. Structural and Thermal Systems of COS
We designed our structural system for the Halo orbiting satellites [also known as Earth Ceres
Communication Orbiters (ECCO)] as a bus that contains all of the instruments and equipment
needed for ECCO. A coilable boom structure is for the satellite array and various
communications systems [1]. A solar array structure is necessary to maintain the shape and
rigidity of the solar panels. The thermal control system for ECCO is a passive thermal control
system (PTCS) that maintains a relatively constant temperature for the instruments inside of the
bus.
The structure that makes the bus is an inch thick honeycomb aluminum structure for
reliability during initial launch compressive load of 6 g’s. Aluminum is the metal used for the
bus specifically for its light weight and strength. Making the aluminum into a honeycomb
structure further increases its strength while decreasing the weight of the overall structure of the
satellite [2] Our satellite is controlled by control moment gyroscopes and reaction wheels, as
opposed to a spin stabilized satellite that is cylindrical, which employs a boxed shaped satellite.
Table 5.9.1.5-1 Total mass and volume of the satellite bus
Mass, kg Volume, m3
Bus Structure 956.9 193.5
There is also a boom structure that connects the main bus to the satellite solar arrays, which
makes use of coilable boom technology. Since the satellite solar array is a massive structure and
may interrupt some of the instruments’ ability to send signals to other communications
equipment, a coilable boom is needed to keep it far away from the bus. Our design for the boom
is interpolated from historical data, in turn providing a mass and volume of the coilable boom:
Detailed Vehicle Descriptions Communications Network Page 528
Author: Leonard Jackson
Table 5.9.1.5-2 Total mass and volume of all of the booms for ECCO
Mass, kg Volume, m3
Coilable Boom Structure 152.2 0.49
We chose coilable boom technology over other ways of deployment attributable to its
compact storage, extreme light weight, and low cost. The material that makes up the coilable
boom is carbon fiber, which enables the boom to flex when stowed, and retain its strength and
stiffness when deployed [ 1].
Figure 5.9.1.5-1 A coilable boom structure segment consisting of longerons (black),
diagonals (red), battens (green), and small joint connectors (blue). This is what the
structure will look like when deployed. The longerons form a triangular cross-section,
which are held together by the battens. The diagonals keep the boom straight by providing
shear stiffness.
ECCO’s solar array panels are based off the design used for the Orion Crew Exploration
Vehicle (CEV): an ultraflex solar array panel [4]. This solar array panel is able to be folded up in
an accordion manner, and when deployed the solar array unfolds into a circular solar array. The
values in the table below include the solar cells as well as the structure:
Table 5.9.1.5-3 Ultraflex solar array mass and volume
Mass, kg Volume, m3
Solar Array Structure 709.6 8.44
By: Leonard Jackson
Detailed Vehicle Descriptions Communications Network Page 529
Author: Leonard Jackson
The ultraflex solar array panel is based on its specific power efficiency; a power to weight
ratio. Given our power requirements for ECCO, we can interpolate data from known systems and
incorporate that data for our system to get a weight for the solar array structure (listed in Table
5.9.1.5-3).
Figure 5.9.1.5-2 The Solar array and structure based off of the Orion CEV. (1) shows the
array when stowed, and is kept at that position until a latch opens. (2) and (3) show the
array unfolding in a circular accordion fashion. (4) shows the array in its final orientation
[3].
Next in the satellite design is the PTCS, which includes reflective coatings such as multi-layer
insulation (MLI), paint, surface finishes, and radiator panels. The reflective coating is much like
the MLI used for the ISPP tanks. The only difference between the tank MLI and satellite MLI is
the number of layers needed. Our satellite uses 15 layers of MLI along with radiator panels to
reject heat into space. An important feature of the ECCO is the color paint they will be using
where the MLI is not able to cover on the satellite. A white paint is used on the surface for the
fact that it has a very low heat absorbance, and high thermal emittance.
We then needed to size our radiator mass and volumes, which is further explained in the
appendix portion. After using a Matlab script, we were able to get a radiator mass and volume
for the ECCO:
Table 5.9.1.5-4 Radiator Panels values for ECCO
Mass, kg Volume, m3
Ceres to Crew Transfer Vehicle 7856.3 2.9
Ceres to Earth 1058.2 2.5
By: NGU Solar Array Technical Paper
Detailed Vehicle Descriptions Communications Network Page 530
Author: Leonard Jackson
References
[1] Unknown Author, “SAILMAST >> What’s It Made of and How Does it Work”
[http://nmp.nasa.gov/st8/tech/sailmast_tech3.html. Accessed Mar. 2011]
[2] Garino, B., Lanphear, J., “Spacecraft Design, Structure, and Operations”, USAF TP, Apr.
2008.
[3] Spence, B., White, S., Wilder, N., Gregory, T., Douglass, M., Takeda, R., “Next
Generation Ultraflex Solar Array for NASA’s New Millennium Program Space
Technology 8”, NASA TP, Dec. 2004.
Detailed Vehicle Descriptions Communications Network Page 531
Author: Tony D’Mello
5.9.1.6. Communications System
5.9.1.6.1 Radio Frequency (RF) Communication
We can see the HOS transmitter dish characteristics in Table 5.9.1.6.1-1. The differences in
each transmitting dish correspond to variations in data rates and frequencies. The receiver dish
characteristics are in Table 5.9.1.6.1-2. Since each rover transmission is in a different frequency,
we require only one receiver which can extract all three signals separately. The second receiver
is for communication between the two satellites because it decreases the pointing error which
results in lower mass, power, and volume requirements.
Table 5.9.1.6.1-1 Transmitter data for one HOS
Rover 1 Rover 2 Rescue Rover HOS
Frequency, GHz 11 7 17 7
Data Rate, HDTV channels 4 4 6 9
Mass, kg 6.76 16.35 6.73 8.89
Power, kW 0.35 0.35 0.1 9
Dish Diameter, m 1.29 1.6 1.6 1.18
Table 5.9.1.6.1-2 Receiver data for one HOS
Rovers HOS
Frequency, GHz 40 7
Data Rate, HDTV channels 7 9
Mass, kg 10.13 8.89
Dish Diameter, m 3.01 1.18
Detailed Vehicle Descriptions Communications Network Page 532
Author: Justin Axsom
5.9.1.6.2 Optical Communication
The Halo Orbiting Satellites (HOS) relay all data from Ceres based operations to Earth or the
Earth-Trailing Relay Satellite (ETRS) through an optical communication system. The details of
the optical communication system appear in Appendix D, Optical Communication Design. We
use the HOS to transmit nine HDTV signals total, corresponding to six crew members and three
cameras. In addition, we also use the HOS to communicate directly with the CTV once it reaches
the half-way point between Earth and Ceres. The CTV continues to communicate with the HOS
through the optical link for the duration of its stay on Ceres. Then, when the crew departs on the
CTV for the return trip, the CTV again communicates with the HOS until it is half-way through
the return transfer to Earth. Each of the HOS have the same optical design and the parameters of
the system appear in Table 5.9.1.6.2-1.
Table 5.4.11-1 Design Parameters of the near Earth Communication Link
Property Value
Wavelength, nm 1064
Data Rate, HDTV signals 9
Earth/ETRS Link CTV Link
Power, kW 37.0 9.27
Mass, kg 124 81.5
Diameter (Rec/Trans), m 4.00 / 0.20 2.00 / 0.40
Length (Rec/Trans), m 1.15 / 0.50 0.76 / 0.50
Detailed Vehicle Descriptions Communications Network Page 533
Author: Graham Johnson
Co-Author: Elle Stephan
5.9.1.7 End of Life Configuration
Once all Ceres operations are complete and the astronauts/CTV return to LEO from Ceres, the
orbits of ECCO 1 and 2 will slowly decay over many years, due to external perturbations, until
all corrective propellant is depleted. Once this occurs, the satellites will eventually orbit with
Ceres about the sun and become part of the asteroid belt as space garbage, but the satellites could
be used as relay satellites for future missions, which delve deeper into the solar system until this
happens.
Detailed Vehicle Descriptions Communications Network Page 534
Author: Kyle Svejcar
5.9.2 Earth-Trailing Relay Satellite (ETRS)
The Earth-Trailing Relay Satellite (ETRS) is a satellite which we place in Earth’s orbit about
the Sun, but trailing the Earth by 90°. This satellite will be used when direct communication
between Ceres and the Earth Communication Network is not possible.
5.9.2.1. Launch Vehicle
When we launch the relay satellite in low Earth orbit (LEO), we require a rocket launch
vehicle to be able to hold the Relay Satellite’s mass and volume. Mass is the real constraint due
to the Relay Satellite having a mass of 37713.9 kg. This larger mass makes the relay satellite too
massive for any rocket launch vehicle except for the Ares V.
Launching into LEO
We use two stages to place the Ares V in a low Earth orbit. The first stage employs 5 Space
Shuttle main engines (SSME), as well as two solid rocket boosters. The propellant for the
SSMEs consists of a liquid hydrogen (LH2) and liquid oxygen (LOX) mixture of propellant.
The thrust is 15,480 kN with a burn time of 110 s.
The second stage consists of 1 J-2X engine, which also employs a propellant of LH2 and
LOX. The thrust for the second stage is 1300 kN and it has a burn time of 465 s.
Detailed Vehicle Descriptions Communications Network Page 535
Author: Megan Sanders
Co-Author: Graham Johnson, Sonia Teran
5.9.2.2. Trajectory
We place the Earth-Trailing Relay Satellite (ETRS) ninety degrees behind Earth in a sun-
centered orbit. This orbit around the Sun is the same as Earth’s, just ninety degrees out of phase.
We do this to ensure that the Sun does not cause any communication loss between Earth and the
astronauts. Keeping the satellite a quarter revolution behind Earth guarantees that even if the
Sun lines up between Earth and Ceres (preventing the Ceres orbiting satellites from connecting
with the main network) a connection is still be possible with the ETRS. This phase difference is
necessary to achieve our mission goal of maintaining uninterrupted HDTV communication with
the astronauts during all parts of the mission.
To place the ETRS at the desired distance away from Earth we employ a transfer orbit which
intersects Earth’s orbit around the Sun at one point. At this location, the satellite performs the
delta-v (ΔV) maneuvers required to move onto and off of the transfer orbit. The intersection
point occurs at the point on Earth’s orbit when the Earth is closest to the Sun, known as
perihelion. We choose perihelion to achieve a tangential burn, which will keep our ΔV
requirements low. The use of perihelion as the intersection point also takes advantage of the fact
that the satellite is traveling the fastest at this point. Though Earth’s orbit in Fig 5.9.2.2-1
appears to be circular, it is actually elliptical due to a small eccentricity of 0.0164. Due to this
slight eccentricity we must perform our ΔV at perihelion or the provided results are no longer
valid and larger ΔV maneuvers are necessary.
Detailed Vehicle Descriptions Communications Network Page 536
Author: Megan Sanders
Co-Author: Graham Johnson, Sonia Teran
Figure 5.9.2.2-1 ETRS transfer orbit shown with Earth orbit. The Earth moves a full
revolution plus 90° degrees while satellite on transfer orbit.
Placing the satellite into position is a multi-step process that calls for the engine to burn at
three different times. We first must have the satellite move from a low-earth holding orbit
around earth to moving alongside the earth around the sun. We accomplish this by having the
satellite take a hyperbolic trajectory to escape the influence of earth. The engines will burn for
the first time to move the satellite off of its low-earth holding orbit and onto the hyperbolic
escape trajectory. The first burn requires a ΔV of 17.371 km/s, the largest of all three steps.
-2.5 -2 -1.5 -1 -0.5 0 0.5 1 1.5 2 2.5
x 108
-2.5
-2
-1.5
-1
-0.5
0
0.5
1
1.5
2
2.5x 10
8
Distance (km)
Dis
tance (
km
)
Earth Orbit
Transfer Orbit
Sun
Earth and Satellite Starting Point, Satellite Ending Point
Earth Ending Point
By Megan Sanders
Detailed Vehicle Descriptions Communications Network Page 537
Author: Megan Sanders
Co-Author: Graham Johnson, Sonia Teran
When the ETRS moves with Earth around the Sun we burn the engine a second time to
move the satellite onto the transfer orbit. This burn requires a delta-v of 1.957 km/s that will be
applied in the same direction as the current velocity, serving to speed up the satellite. We select
this particular transfer orbit because it has a period of 1.25 years, meaning that in the time it
takes the satellite to circle the Sun once the Earth will have circled the Sun 1.25 times. Because
the Earth moves further around the Sun, it is ninety degrees ahead of the satellite, completing our
goal.
We burn the engine for the final time when the satellite has returned to its starting point at the
intersection between the transfer orbit and Earth’s orbit. This burn will also require a delta-v of
1.957 km/s but it will be applied in the direction opposite the current velocity, serving to slow
down the satellite. At the completion of the burn the ETRS is on Earth’s orbit around the sun,
but ninety degrees behind Earth.
Table 5.9.2.2-1 ΔV requirements for Earth-Trailing Relay Satellite
ΔV requirement (km/s)
Escaping Earth 17.371
Entering Transfer Orbit 1.957
Leaving Transfer Orbit 1.957
Total 21.285
The total ΔV we need for all of these maneuvers is 21.285 km/s, with the majority coming
from the escape from Earth. After the completion of these maneuvers the relay satellite remains
behind Earth in Earth’s orbit around the Sun until the end of the mission.
A more detailed discussion of the thought process behind this configuration and the algorithm
used to calculate it can be found in Section A.5.9.2.2.
Detailed Vehicle Descriptions Communications Network Page 538
Author: Elle Stephan
5.9.2.3. Power Systems
The ETRS power system implements the same photovoltaic cells as the Ceres orbiting
satellites. The main difference being that the ETRS is in the same orbital plane as Earth,
allowing for nine times the solar irradiation of 1353W/m2 [1]. Taking into account the
photovoltaic cell efficiency of 42.3%, the power generated by the ETRS is 572.3W/m2 [2].
Given the power required to for operation of 52kW, the array outputs 54.9kW of power. In
order to generate this power, two circular arrays with eight panels each have a combined area of
96 sq. meters and combined mass of 80.64kg. Each panel is an isosceles triangle with a height of
4m and a length of 3m. When fully deployed, each array will have a diameter of 8m.
Figure 5.9.2.3 - 1 The solar arrays (in yellow) provide the power necessary to operate the
ETRS.
For the duration of the mission, the solar array should be in constant sight of sunlight.
However, if something should happen and the array is blocked from direct sunlight, a Li-ion
rechargeable battery compensates by providing the necessary power.
Detailed Vehicle Descriptions Communications Network Page 539
Author: Kyle Svejcar
5.9.2.4. Propulsion Systems
We place the relay satellite into its orbit with a three burn sequence. This sequence consists
of one burn to leave Earth’s orbit, one burn to enter the transfer orbit, and one burn to leave the
transfer orbit. The Earth-Trailing Relay Satellite takes 1.25 years to transfer into the Earth
trailing orbit.
For the ETRS, we design two engines to generate a total thrust of 50398 N, with a
corresponding Isp of 460s at vacuum conditions. Our system expends a bipropellant mixture with
liquid hydrogen (LH2) as the fuel and liquid oxygen (LOX) as the oxidizer; we use this mixture
at an oxidizer to fuel ratio of 5.22:1. Our nozzle has an expansion ratio of 150.9.
We keep the nozzle cool using regenerative cooling. Regenerative cooling works by running
cold propellant through a heat exchanger. The propellant absorbs heat being transferred to the
structure, allowing the structure to maintain a lower temperature [1].
The fuel feeds from the tanks to the nozzle using a pump system due to the tanks being
greater than ten m3. We use turbopumps to pump the propellant from the propellant tanks to the
combustion chamber.
Detailed Vehicle Descriptions Communications Network Page 540
Author: Kyle Svejcar
5.9.2.4.1 Inert Mass and Volume Breakdown
The fuel and oxidizer tanks are made out of carbon fiber composite. Table 5.9.2.4.1-1 shows
the breakdown of our masses and volumes for ECCO base.
Table 5.9.2.4.1-1 Mass and volume breakdown for ECCO Base
Component Mass, kg Volume, m3
LH2 5116.8 --
LOX 26,300 --
2 Nozzles 233.0 31.7
2 LH2 tanks 60.0 72.0
2 LOX tanks 37.6 23.0
Pump system 29.5 --
Support Structure 33.1 --
5.9.2.4.2 Nozzle Dimensions
Table 5.9.2.4.2-1 gives the dimensions for ERTS’s nozzle based on the dimensions in Fig.
5.9.2.3-1.
Figure 5.9.2.4.2-1 General Shape of Nozzle
Variable Length, mm Angle, deg
Rc 84.56 --
R2 171.7 --
Lc 205.3 --
Rt 32.105 --
Rn 12.26 --
Le 1082.88 --
Re 394.365 --
Rl 48.16 --
b -- 30
Tn -- 39
Te -- 8
Detailed Vehicle Descriptions Communications Network Page 541
Author: Kyle Svejcar
5.9.2.4.3 Conclusion
The total mass for our propulsion system for the relay satellite is 31776.9 kg and the volume
is 126.7 m3.
Detailed Vehicle Descriptions Communications Network Page 542
Author: Kyle Svejcar
References
[1] Humble, Ronald W., Gary N. Nelson, and Wiley J. Larson. Space Propulsion Analysis and
Design. 1st Ed., Revised. McGraw-Hill, 1995.
Detailed Vehicle Descriptions Communications Network Page 543
Author: David Schafer
5.9.2.5. Attitude Determination and Control
The Relay Satellite needs to be near perfectly stabilized in order to transmit the optical signal
from Ceres to Earth. This means that the attitude determination system and the actuators
together need to be accurate to about 0.02 degrees (400 microradians). Additionally, this
satellite must provide continuous communications from Earth to Ceres and back. This means no
blackout periods are allowed, making the satellite perfectly stable for all time.
5.9.2.5.1 Attitude Positioning
Once the Relay Satellite has reached its orbit, the only positional stability issues become
correcting for the small environmental forces in Earth’s orbit of the Sun. We achieve the
knowledge on the specific location of the satellite through the use of Motion Reference Units
(MRU), an inertia referencing system that uses actuators to find the forces acting on the body. A
single MRU triplet is required, finding accelerations in all directions. This information needs to
be integrated to find velocities and the ultimate position of the satellites, so an onboard computer
needs to be coupled with this system in order to find exact values of position and velocity at any
given point in time, and to map out where exactly the satellite is, yet the requirement of tracking
all the other vehicles and structures already calls for a computer, so implementing this MRU
system adds no extra mass other than the MRU itself.
Over-estimating the forces acting on the Relay satellite, we arrive at a total of about 0.0014 N
acting on the satellite at any one point in time. This force acts on the satellite in any direction,
meaning the satellite must be able to correct from any direction. We use the sole thruster on the
satellite to counteract the force. This means that the single thruster must rotate to counteract the
forces in the proper direction. This rotation is achieved by rotating the full satellite when not in
Detailed Vehicle Descriptions Communications Network Page 544
Author: David Schafer
use, and firing the thruster to bring the Relay Satellite back to the desired position. Obviously,
this can’t be done continuously, as then the satellite would never be in use. Additionally, the
nature of the difference between the thruster force and the environmental forces shows that a
continuous thrusting would create far too high of a corrective thrust, requiring more and more
corrective thrusts. Despite the intermittent thrusting, table 5.9.2.5.2-1 shows that the propellant
mass is considerable. This large mass is due to the 7-year mission life of the satellite.
5.9.2.5.2 Attitude Pointing
We can split our satellite pointing control issues into two main problems: sensing direction
with great accuracy and changing pointing direction while maintaining this accuracy. Each
aspect introduces its own problems, yet they are all easily solvable.
Attitude Determination
The attitude determination system of the Relay Satellite is the same system implemented by
the Halo Orbiting Satellites presented in section 5.9.1.4.
Actuation
The Relay Satellite implements a dual system of Control Moment Gyroscopes (CMG) and
Reaction Wheels in order to fulfill the positioning requirements. Both systems are very accurate,
with the CMGs having the lower accuracy of about 350 microradians [2]. The CMGs are the
main system, but as they approach saturation the reaction wheels start up and control the satellite
as the CMGs pass through saturation. Once the system has fully passed through saturation, the
reaction wheels shut down again and let the CMGs take full control of the satellite back over. As
the torques to be controlled are very small in magnitude – only about 0.02 Newton-meters of
torque – both the CMGs and the reaction wheels are quite small, requiring very little volume,
Detailed Vehicle Descriptions Communications Network Page 545
Author: David Schafer
and, most importantly, have very small power requirements. The full attitude control system
values are shown below in table 5.9.2.5.2-1
Table 5.9.1.4.2-1 Full mass, power, and volume requirements of the entire attitude control
system on the Relay Satellite
Hardware Mass, kg Power, kW Volume, m3
MRU 2.5 0.012 0.01
Fine Guidance Sensor 660 0.06 1.275
CMGs 20 0.8 0.0062
Propulsion System 13.28 0 0.8
Propellant 45 0 0.005
Saturation effects
The need for the dual system arises from the tendency of both the CMGs and the reaction
wheels to reach a saturation point – a point where they can no longer provide any torque to the
satellite. Using simple computer logic, the CMGs can perturb a single gyro and torque against
that until that portion of the dual system has passed through its saturation point [1]. Then, the
perturbed gyro is released and left to correct against the torques it was earlier. This means the
satellite passes by the saturation point without losing control. Yet this means the satellite itself
loses connection to both Earth and Ceres. In order to maintain this connection, the reaction
wheels torque the satellite to keep connection with the rest of the satellite network. This system
uses the CMGs as the primary actuating system, keeping the Satellite pointed properly for most
of the time, and the reaction wheels only become active when the CMGs are near their saturation
points.
Detailed Vehicle Descriptions Communications Network Page 546
Author: David Schafer
References
[1] Bedrossion, Nazareth. Bhatt, Sagar. Alaniz, Abrin. McCantz, Edward. Nguyen, Louis. and
Chamitoff, Greg. “ISS Contingency Attitude Control Recovery Method For Loss Of
Automatic Thruster Control” American Astronautics Society. February 1-6, 2008
[2] Bradford Engineering “Reaction Wheel Unit” Space Systems and Components Division.
Model W18. September, 2006
Detailed Vehicle Descriptions Communications Network Page 547
Author: Leonard Jackson
5.9.2.6. Structural and Thermal Systems
The structural system for the Earth Trailing Relay Satellite (ERTS), named ECCO Base, is
very similar to the ECCO 1 and 2 satellites in section 5.9.1.5. The structure consists of a bus, a
coilable boom, and a satellite array structure. The thermal control system is still a PTCS just like
ECCO 1 and 2.
The bus for the ERTS is also made from a 1 inch aluminum honeycomb structure, and is
controlled the same way the halo satellites are controlled. There are slight differences in ERTS’s
dimensions as it is closer to Earth and the Sun. The table below is a summary of the bus
structure.
Table 5.9.2.6-1 Total mass and volume of the satellite bus
Mass, kg Volume, m3
Bus Structure 1364.1 270.8
The boom structure for the satellite array has correspondingly changed due to the lower area
requirement of the solar arrays. It still employs a coilable boom to deploy the solar array [1].
Table 5.9.2.6-2: Total mass and volume of all of the booms for ETRS
Mass, kg Volume, m3
Coilable Boom Structure 23.6 0.01
The solar array panels on ERTS also make use of the Orion CEV ultraflex solar array panels,
and are smaller than those found on the other communication satellite solar arrays because ETRS
receives more solar energy [2].
Table 5.9.2.6-3: Ultraflex solar array mass and volume
Mass, kg Volume, m3
Solar Array Structure 80.6 0.96
Detailed Vehicle Descriptions Communications Network Page 548
Author: Leonard Jackson
The PTCS of ERTS include the same components as the other communication satellites. The
radiator panels were also designed with our radiator sizing Matlab script.
Table 5.9.2.6-4: Radiator panel parameters for ETRS
Mass, kg Volume, m3
Satellite to Crew Transfer Vehicle 81.5 5.5
Satellite to ECCO 124.0 12.1
Satellite to Earth 72.1 0.9
ETRS also includes extra radiator paneling because it needs an extra communication link,
which leads to more heat generation within ETRS. For the most part, ETRS is a modified ECCO
satellite; for more detailed descriptions, refer back to Section 5.9.1.5.
Detailed Vehicle Descriptions Communications Network Page 549
Author: Leonard Jackson
References
[1] Unknown Author, “SAILMAST >> What’s It Made of and How Does it Work”
[http://nmp.nasa.gov/st8/tech/sailmast_tech3.html. Accessed Mar. 2011]
[2] Spence, B., White, S., Wilder, N., Gregory, T., Douglass, M., Takeda, R., “Next
Generation Ultraflex Solar Array for NASA’s New Millennium Program Space
Technology 8”, NASA TP, Dec. 2004.
Detailed Vehicle Descriptions Communications Network Page 550
Author: Justin Axom
Co-Author: Tony D’Mello
5.9.2.7. Communications System
The Earth-Trailing Relay Satellite (ETRS) has three optical communication modules. Since
we use the ETRS to communicate with Earth when we cannot establish a direct link with Earth,
it needs to communicate simultaneously with the Ceres Halo Orbiting Satellites (HOS), crew
transport vehicle (CTV), and the Earth Orbiting Satellites (EOS). We accomplish this task by
using a separate optical link for each. This way, the modules independently aim at each vehicle
in order to complete the communication link. Refer to Appendix D for a more in-depth
discussion of the optical communication system and appendix A-5.9.2.7 for more information on
the ETRS design considerations. The design parameters for each module appear in table 5.9.2.7-
1.
Table 5.9.2.7-2 Design Parameters of the near Earth Communication Link
Property Value
Wavelength, nm 1064
Data Rate, HDTV signals 9
EOS Link CTV Link HOS Link
Power, kW 5.30 9.27 37.0
Mass, kg 72.1 81.5 124
Diameter (Rec/Trans), m 1.50 / 0.50 2.00 / 0.40 4.00 / 0.20
Length (Rec/Trans), m 0.38 / 0.50 0.76 / 0.50 1.15 / 0.50
Detailed Vehicle Descriptions Communications Network Page 551
Author: Elle Stephan
5.9.2.8. End of Life Configuration
After the mission is completed and the astronauts are safely returned to Earth, the ETRS will
remain in orbit around the Sun. It can still act as a relay satellite for future missions. The
attitude control thrusters will eventually deplete their fuel supply, and at this point in time,
communication will no longer be reliable.
Detailed Vehicle Descriptions Communications Network Page 552
Author: Justin Axsom
5.9.3 Earth-Orbiting Receiving Satellites
We use the Earth-Orbiting Satellites (EOS) to transmit and receive all optical signals to and
from Earth. We equipped the EOS satellites with a radio frequency (RF) system as well as an
optical system. The RF system transmits and receives all signals from Earth. Once the EOS
receives the RF signal from Earth, the computer aboard transcodes the signal to an optical signal.
The optical system then transmits the signal to either the Ceres orbiting halo satellites, Earth
trailing relay satellite, or the crew transport vehicle. When receiving a signal, the process works
in reverse where an optical signal is transcoded into and RF signal, then relays to Earth. Refer to
appendix D.3.1.1 for a more in-depth discussion of the optical system. We plan to incorporate
the EOS into the NASA Tracking and Data Relay Satellite System to handle deep space optical
communication, and assume that the EOS will be launched and in place prior to our mission.
Refer to appendix A-3.2.9.1.1 for more information on the ground based operations and
appendix A-5.9.3 for the design considerations behind the EOS.
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