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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


5.4.3. Outbound Trajectory .....................................................................................181

5.4.4. Power Systems ...............................................................................................187



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


5.5. Crew Capsule ..................................................................................................................329


5.5.2. Power Systems ...............................................................................................331




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


5.6.2. 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


5.7.2. Power Systems ...............................................................................................421



5.7.5. Attitude Determination and Control Systems ................................................435

5.7.6. Structural and Thermal Control Systems ......................................................439


5.7.8. Ceres Rock Collection Process ......................................................................457

5.7.9. Science Toolbox and Experimentation ..........................................................458

5.7.10. Autonomous Operations ..............................................................................459


5.8. Rescue Rover ..................................................................................................................465


5.8.2. Power Systems ...............................................................................................466

AAE 450 – Spacecraft Senior Design – Spring 2011 Project Vision - v



5.8.5. Attitude Determination and Control Systems ................................................485

5.8.6. Structural and Thermal Control Systems ......................................................439


5.8.8. Trajectory and Flight Path .............................................................................501

5.8.9. Autonomous Operations ................................................................................503


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


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


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.

https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2011/spring


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.


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.


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


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%.


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.


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?


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.


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?


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.


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


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


“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




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


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




ISPP Stations

Table 3.1-6 Nominal technical specifications for one of the identical ISPP stations


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)




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


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



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



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


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.


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


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


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.


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.


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



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.



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].

http://www.nasa.gov/mission_pages/constellation/ares/aresV/index.html

http://www.nasa.gov/pdf/366590main_Ares_I_FS.pdf


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



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



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.



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.



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.


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.



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



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.


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



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.



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



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.


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


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.


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.


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.



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.



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



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.



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.


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.


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.



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



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


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.




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




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].




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


Project Conclusions Page 67

4. Project Conclusions

Project Conclusions General Project Concerns Page 68


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.



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.



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


Author: David Wyant


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


Author: David Wyant


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


Author: David Wyant


assure that the water is fully extracted from the regolith, another mission parameter requires

heating the regolith to 200oC above zero.


Author: David Wyant


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.



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].



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



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



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.



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


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.




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.




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.




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


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.



4.3.1 Vehicle Costs


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.



Table 4.3.1-2: STV Advanced Mission Cost Model breakdown



Dry Mass kg 62,000

Specification - - Planetary


Block - - 1

Difficulty - - High

Cost


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



Dry Mass kg 28,000

Specification - - Planetary Lander


Block - - 1

Difficulty - - Very High

Cost


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



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



Dry Mass kg 11,510

Specification - - Rover


Block - - 2

Difficulty - - Average

Cost


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



Dry Mass kg 6,200

Specification - - Rover


Block - - 1

Difficulty - - High

Cost


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



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



Dry Mass kg 13,000

Specification - - Communication


Block - - 2


Cost


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



Dry Mass kg 8,800

Specification - - Communication


Block - - 2

Difficulty - - Low

Cost




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



Dry Mass kg 10,000

Specification - - Manned Re-entry


Block - - 2


Cost




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




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



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.



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.



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


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.



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.



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.



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.



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

http://www.nasa.gov/mission_pages/constellation/ares/aresl/index.html



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


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


8 LOX Tank 5 146920.00 181.29


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


15 LOX Tank 3 145850.00 181.29

16 LOX Tank 4 145850.00 181.29


17 LOX Tank 5 145850.00 181.29


18 Reactor 22926.27 296.00

19 Comm. Satellites 28347.81 463.53




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 --




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




launches 1 and 10. These tanks will be attached so that they can be separated before the supply

transfer vehicle departs.



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



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.



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.



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


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.


Author: Sonia Teran

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


Author: Sonia Teran

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


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


Author: Sonia Teran

known systems chemical rockets were chosen. A more detailed discussion can be found in

Appendix H.


Author: Sonia Teran

References

[1] “HORIZONS Web Interface,” Solar System Dynamics, URL:

http://ssd.jpl.nasa.gov/horizons.cgi#results [cited 26 March 2011].


Author: Alex Park

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.


Author: Alex Park

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


Author: Alex Park

Figure 5.2.4-2 Model of the STV power system


Author: Alex Park

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


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


Author: Alex Park

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


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


Author: Alex Park

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.


Author: Alex Park

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.


Author: Alex Park

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


Author: Alex Park

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.


Author: Alex Park

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.


Author: Alex Park

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



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)



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.


Author: Alex Park

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.


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



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



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



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.



Fig. 5.2.5.4-1 CRE nozzle geometry from RPA

By: 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




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


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


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.


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


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


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


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.


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.



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



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



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



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



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



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



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


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.


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.



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.




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.



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



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



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].

http://deepspace.jpl.nasa.gov/advmiss/docs/NASA_MO&CS.pdf



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

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2

3

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Evan Helmeid

Trajectory of Spacecraft

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Y-p

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Final trajectory

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Evan Helmeid

Steering Angle, , vs Time

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Ste

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Evan Helmeid

Y-Velocity vs X-Velocity

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Evan Helmeid

Velocity Components vs Time

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Velo

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Evan Helmeid


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Evan Helmeid


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By: 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.



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.


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.

Detailed Vehicle Descriptions Crew Launch Vehicle Page 167


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



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.



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



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.



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]

http://www.nasa.gov/pdf/156298main_orion_handout.pdf

http://www.nasa.gov/mission_pages/constellation/ares/aresl/index.html

http://www.orbital.com/NewsInfo/Publications/Las_Fact.pdf

Detailed Vehicle Descriptions Crew Transfer Vehicle Page 173

Author: Trey Fortunato, Chris Luken

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



5 Ares V Earth Departure Tank 1



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



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



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



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



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.



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



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



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



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.



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



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.



(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



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.



References

[1] ―HORIZONS Web Interface,‖ Solar System Dynamics, URL:

http://ssd.jpl.nasa.gov/horizons.cgi#results [cited 26 March 2011].


Author: Drew Crenwelge

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



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



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



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:



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



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



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



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



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.


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



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)


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



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



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.



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.



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.)


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.




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




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


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


Author: Alex Kreul


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.


Author: Alex Kreul


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


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


Author: Alex Kreul


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.


Author: Alex Kreul


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


Author: Alex Kreul


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 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




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.




Table 5.4.5.3-, and Table 5.4.5.3-.




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




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




Volume, m3 2.09

Mass, kg 23.86


Author: Alex Kreul


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.


Author: Alex Kreul


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




Volume of 1, m3 0.22


Mass of 1, kg 5.27

Mass of 6, kg 31.59


Author: Alex Kreul


Table 5.4.5.4-2 Attitude Thruster Tank System specifications. This table includes the LOX

tank dimensions, mass, and volume.

Sphere






Mass of 1, kg 4.40

Mass of 6, kg 26.39


Author: Alex Kreul


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


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


Author: Chris Luken


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.


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.


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


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


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


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


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.


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.


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


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



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.



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.



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



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:



(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



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



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



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



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



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



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


IonR

ail

Y P

ositio

n (

m)

MPD Thrusters

Max Radius

Burn Arc Boundary



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)



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


IonR

ail

Y P

ositio

n (

m)

MPD Thrusters

Max Radius


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



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



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.



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.



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



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.



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:



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



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


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].



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.



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.



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.



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.




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




References

[1] Hanford, A., ―Advanced Life Support Baseline Values and Assumptions Document,‖

NASA/CR—2004—208941, Aug. 2004, p. 22



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.



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



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



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.



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


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.



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



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


Crew Transfer Vehicle 334.8 2.790 9.30

Drawing:Jillian Roberts

Photo: Daniel Shultz



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.



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).



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.



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.



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.



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.



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



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.


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.



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.



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.



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.



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.



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



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.



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



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]



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



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.



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



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


Radiation Dosimeters 0.1345 0 0.003



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.

http://sragnt.jsc.nasa.gov/Publications/TM104782/techmemo.htm.%20Jan%202011

http://spaceflight.nasa.gov/shuttle/reference/shutref/crew/radiation.html


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


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.


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



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


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.


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.


Author: Kim Madden

Table 5.4.9.-1 CTV thermal control system summary

Component Mass , kg Power , kW Volume, m3


Heat Pipe 333.86 0 3.76

Radiators 4,237.62 1.72 1.51


Heater 71.36 0 0.14

TOTAL 4,729.32 1.72 5.62


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



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



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®


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



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


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



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


Max Allowable Temperature 650 oC

Material Density 1500 kg/m3

Tensile Strength 5650 MPa



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



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


Number 4 --

Tether Radius 0.0325 m


Tether Volume 0.4446 m3




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.


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)



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.



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



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



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].

http://tpsx.arc.nasa.gov/

Detailed Vehicle Descriptions Crew Transfer Vehicle Page



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



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



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.



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



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



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.


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.



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

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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


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


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.



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



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


Wet mass 184.6 T

Dry mass 160.6 T


Thrust – nominal 10~100 kN

Tnominal:W range 2.0~2.3 - -

Total transfer time 85.66 min



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


Wet mass 471.4 T

Dry mass 424.2 T


Thrust @ 30% 450.0 kN

T30%:W range 3.5~3.9 - -

Burn time 468.3 s



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)



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.



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



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.


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.



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.



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



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



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



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



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 Crew Capsule Page 329


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.



Table 5.5.1-2 Mass, power, and volume for the Crew Capsule


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.


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


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].


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]


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.


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


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.

http://www.astronautix.com/craft/orionlas.htm

http://www.orbital.com/NewsInfo/Publications/Industry_Pad_Fact.pdf



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


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 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


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


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


Food System 138.9 0 9.01



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


Air System 30.53 0 9.01


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



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


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.



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.



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



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



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



References

[1] “Constellation Orion Crew Exploration Vehicle”, NASA Fact Sheet No. FS-2008-07-031-

GRC, January 2009.


Author: Kim Madden


The details for the Crew Capsule thermal control system can be found in Section 5.7.11.2.


Crew Capsule thermal control system.

Table 5.5.6.4-1: Crew Capsule Thermal Control System Summary


Heat Pipe 6.85 0 0.98

Radiators 53.79 0 0.05


TOTAL 67.67 0 1.04


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.



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.



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



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



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.



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


Number 4 --

Tether Radius 0.0175 m


Tether Volume 0.0385 m3




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



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


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



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.



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



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.



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



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



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].

http://tpsx.arc.nasa.gov/




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



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


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.


Author: Paul Frakes

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.


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


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.



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


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.


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.


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.


Author: Paul Frakes

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.


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.



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.



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.



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.



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.



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.



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

Detailed Vehicle Descriptions In-Situ Propellant Production Stations Page 381

Authors: Matthew Hill, Zachary Richardson

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



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



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



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.



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


Author: Matthew Hill

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).



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



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



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.



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



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

By: Alex Roth



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



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



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



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.



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.



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.


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



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.



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.



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.



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.



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.



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.



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



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



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.

http://www.protonenergy.com/



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



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



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


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



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



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>.


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




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.




References

[1] [fomo.com/resources/technical-bulletins/openvsclosed.aspx. Accessed Feb. 2011.]


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.



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



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.



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



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


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.


Author: Joel Lau

Figure 5.7.2-1 A generator converts mechanical power to electrical to satisfy the

Exploration Rovers’ electrical power needs


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.


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


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


Author: David Wyant


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.


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.


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].


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.


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.


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.


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.]

http://www.deere.com/servlet/ProdCatProduct?tM=FR&pNbr=9870SH

http://scienceworld.wolfram.com/physics/CoefficientofRestitution.html



Co-Author: Jillian Roberts


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


Water Supply and

Regeneration

4 7 408.1 0.160 0.608




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.




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.



Lighting System 183.6 1.530 5.100


Author: David Schafer


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.



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



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.



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.


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.


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


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



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


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


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


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


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


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.


Author: Kim Madden


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.


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.



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


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.


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.



Author: Kim Madden

Results and Summary


Exploration Rover thermal control system.

Table 5.7.6.2-1 Exploration Rover Thermal Control System Summary



Heat Pipe 29.876 0 0.302

Radiators 304.038 0.061 0.108


Heater 15.378 0 0.028

TOTAL 367.557 0.061 0.481


Author: Kim Madden

References








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


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



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.



Table 5.7.7.2-2 External communication device characteristics for Rover 2


Frequency, GHz 26.5 7


Mass, kg 0.67 16.35

Power, kW 0.1 __




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.



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.



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.



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].



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.



References

[1] “DARPA Urban Challenge” DARPA Grand Challenge. 2007


Author: Kim Madden


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 Rescue Rover Page 465


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.


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



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.



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



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


Author: David Wyant



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].


Author: David Wyant


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


Author: David Wyant


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.


Author: David Wyant


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


Author: David Wyant


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].


Author: David Wyant


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.


Author: David Wyant


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.]

http://www.launchpnt.com/Documents/Dual%20Halbach-Motor-Data-Sheet_R1.pdf

http://www.launchpnt.com/Documents/Dual%20Halbach-Motor-Data-Sheet_R1.pdf





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


Water Supply and

Regeneration

4 2 244.8 0 0.245




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.




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.



Lighting System 108 0.900 3




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].



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



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



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


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



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.



References

[1] Kongsberg Maritime “MRU 5+ Datasheet,” Trondheim, Norway

[2] Royer, Caleb, “Robotic Canfield Joint,” National Instruments, April 29, 2010


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.


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,


R

C

h

D


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



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


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



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



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


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


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


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.


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.


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



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.


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


Rescue Rover thermal control system.

Table 5.8.6.2-1 Rescue Rover thermal control system summary



Heat Pipe 25.301 0 0.218

Radiators 184.857 0.016 0.066


TOTAL 228.160 0.016 0.325


Author: Kim Madden

References








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


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



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


Frequency, GHz 60 17


Mass, kg 0.37 5.92

Power, kW 0.1 __



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.


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.



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



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.




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


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



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


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.



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.




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.




*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




*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)




*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




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.


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.


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].

http://www.solarfeeds.com/energy-boom/14566-spire-announces-worlds-most-efficient-concentrated-pv-solar-cell

http://www.solarfeeds.com/energy-boom/14566-spire-announces-worlds-most-efficient-concentrated-pv-solar-cell


Author: Kyle Svejcar


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.




Table 5.9.1.3.1-1 Mass and volume breakdown for ECCO 1


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


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 --


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].




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.




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.



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.




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.




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


H2O2 1248 --

H2O2 tank 5.1 0.73

Feed system 4.4 --

Nozzle 0.69 0.06


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


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.




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-




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


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


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.




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).




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


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:



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



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



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.

http://nmp.nasa.gov/st8/tech/sailmast_tech3.html



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


Mass, kg 10.13 8.89




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



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.



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.



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.




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




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.



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.



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.



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


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 --


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


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



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.



References

[1] Humble, Ronald W., Gary N. Nelson, and Wiley J. Larson. Space Propulsion Analysis and

Design. 1st Ed., Revised. McGraw-Hill, 1995.



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



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,



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


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


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.



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



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



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.



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.

http://nmp.nasa.gov/st8/tech/sailmast_tech3.html


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



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.



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.