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Purdue University

AAE 450 Spacecraft Design

Spring 2009

Project

Xpedition

Contents

1 – FOREWORD ........................................................................................................................................... 5

2 – INTRODUCTION ................................................................................................................................... 7

2.1 – BACKGROUND .................................................................................................................................... 7 2.2 – WHAT‟S IN THIS REPORT? .................................................................................................................. 9 2.3 – ACKNOWLEDGEMENTS .....................................................................................................................10 2.4 – ACRONYM LIST .................................................................................................................................12

3 – PROJECT OVERVIEW ....................................................................................................................... 14

3.1 – DESIGN REQUIREMENTS ...................................................................................................................14 3.2 – INTERPRETATION OF DESIGN REQUIREMENTS ...................................................................................18 3.3 – DESIGN PROCESS ..............................................................................................................................19 3.4 – RISK AND COST ANALYSIS ................................................................................................................22

4 – RESULTS ............................................................................................................................................... 25

5 – MISSION CONFIGURATION 100G PAYLOAD ............................................................................. 27

5.1 – SYSTEM OVERVIEW ..........................................................................................................................28 5.2 – LAUNCH VEHICLE .............................................................................................................................31

5.2.1 – Launch Vehicle Selection .........................................................................................................31 5.2.2 – Earth Parking Orbit Selection .................................................................................................33 5.2.3 – Attitude Determination in LEO ................................................................................................35

5.3- LUNAR TRANSFER ..............................................................................................................................37 5.3.1 – Structure ...................................................................................................................................37 5.3.2 – Mission Operations ..................................................................................................................47 5.3.3 – Propulsion on the Orbital Transfer Vehicle .............................................................................51 5.3.4 – Attitude .....................................................................................................................................56 5.3.5 – Communication ........................................................................................................................60 5.3.6 – Thermal Control .......................................................................................................................67 5.3.7 – Power .......................................................................................................................................69

5.4 – LUNAR DESCENT ..............................................................................................................................74 5.4.1 – Structures .................................................................................................................................74 5.4.2 – Mission Operations ..................................................................................................................82 5.4.4 – Attitude Control .......................................................................................................................93 5.4.5 – Communication ........................................................................................................................95 5.4.6 – Thermal Control .......................................................................................................................98 5.4.7 – Power .....................................................................................................................................104

5.5 – LOCOMOTION ..................................................................................................................................109 5.5.1 – Structures/Integration ............................................................................................................110 5.5.2 – Propulsion ..............................................................................................................................115 5.5.3 – Communications ....................................................................................................................117 5.5.4 – Thermal Control .....................................................................................................................120 5.5.5 – Power .....................................................................................................................................121 5.5.6 – Mission Operations ................................................................................................................123

5.6 – MISSION INTEGRATION ...................................................................................................................126 5.7 – RISK ANALYSIS ...............................................................................................................................132 5.8 – COST ANALYSIS ..............................................................................................................................133

6 – MISSION CONFIGURATION 10KG PAYLOAD .......................................................................... 134

6.1 – SYSTEM OVERVIEW ........................................................................................................................135 6.2 – LAUNCH VEHICLE ...........................................................................................................................138

6.2.1 – Launch Vehicle/Site................................................................................................................138 6.2.2 – Earth Parking Orbit Selection ...............................................................................................139

6.2.3 – Attitude determination in LEO ...............................................................................................140 6.3 – LUNAR TRANSFER ...........................................................................................................................141

6.3.1 – Structure .................................................................................................................................141 6.3.2 – Mission Operations ................................................................................................................145 6.3.3 – Propulsion on the Orbital Transfer Vehicle ...........................................................................147 6.3.4 – Attitude ...................................................................................................................................148 6.3.5 – Communication ......................................................................................................................150 6.3.6 – Thermal Control .....................................................................................................................151 6.3.7 – Power .....................................................................................................................................152

6.4 – LUNAR DESCENT ............................................................................................................................155 6.4.1 – Structures ...............................................................................................................................155 6.4.2 – Mission Operations ................................................................................................................163 6.4.3 – Propulsion ..............................................................................................................................165 6.4.4 – Attitude Control .....................................................................................................................171 6.4.5 – Communication ......................................................................................................................172 6.4.6 – Thermal Control .....................................................................................................................173 6.4.7 – Power .....................................................................................................................................175

6.5 – LOCOMOTION ..................................................................................................................................179 6.5.1 – Trajectory ...............................................................................................................................179 6.5.2 – Propulsion ..............................................................................................................................181 6.5.3 – Attitude ...................................................................................................................................183


7 – MISSION CONFIGURATION LARGE PAYLOAD ...................................................................... 188

7.1 – SYSTEM OVERVIEW ........................................................................................................................189 7.2 – LAUNCH VEHICLE ...........................................................................................................................192

7.2.1 – Launch Vehicle/Site................................................................................................................192 7.2.2 – Earth Parking Orbit ...............................................................................................................194 7.2.3 – Attitude determination in LEO ...............................................................................................195

7.3 – LUNAR TRANSFER ...........................................................................................................................196 7.3.1 – Structure .................................................................................................................................196 7.3.2 – Mission Operations ................................................................................................................200 7.3.3 – Propulsion on the Orbital Transfer Vehicle ...........................................................................202 7.3.4 – Attitude ...................................................................................................................................204 7.3.5 – Communication ......................................................................................................................206 7.3.6 – Thermal Control .....................................................................................................................207 7.3.7 – Power .....................................................................................................................................208

7.4 – LUNAR DESCENT ............................................................................................................................211 7.4.1 – Structures ...............................................................................................................................211 7.4.2 – Mission Operations ................................................................................................................213 7.4.3 – Propulsion ..............................................................................................................................214 7.4.4 – Attitude Control .....................................................................................................................218 7.4.5 – Communication ......................................................................................................................219 7.4.6 – Thermal Control .....................................................................................................................220 7.4.7 – Power .....................................................................................................................................222

7.5 – LOCOMOTION ..................................................................................................................................225 7.5.1 – Hover Trajectory ....................................................................................................................225 7.5.2 – Propulsion ..............................................................................................................................227 7.5.3 – Attitude Control .....................................................................................................................228


8 – ALTERNATIVE DESIGNS ............................................................................................................... 232

8.1 – LAUNCH VEHICLE ALTERNATIVE DESIGNS .....................................................................................234 8.1.1 – Launch Vehicle Alternatives ..................................................................................................234

8.2 – LUNAR TRANSFER ALTERNATIVE DESIGNS ....................................................................................236 8.2.1 – Trajectory Alternatives ..........................................................................................................236 8.2.2 – Propulsion Alternatives ..........................................................................................................242 8.2.3 – Attitude Alternatives ...............................................................................................................246 8.2.3 – Power Alternatives .................................................................................................................247

8.3 – LUNAR DESCENT ALTERNATIVE DESIGNS ......................................................................................251 8.3.1 – Landing Alternatives ..............................................................................................................251 8.3.2 – Structural Alternatives ...........................................................................................................266 8.3.3 – Trajectory Alternatives ..........................................................................................................267 8.3.4 – Propulsion Alternatives ..........................................................................................................269 8.3.5 – Attitude Alternatives ...............................................................................................................272 8.3.6 – Thermal Control Alternatives ................................................................................................276 8.3.7 – Power Alternatives .................................................................................................................277

8.4 – ROVER ............................................................................................................................................278 8.4.1 – CAD/Integration ....................................................................................................................280 8.4.2 – Communications ....................................................................................................................281 8.4.3 – Propulsion ..............................................................................................................................282 8.4.4 –Thermal Control ......................................................................................................................284 8.4.5 – Power .....................................................................................................................................285 8.4.6 – Deployment ............................................................................................................................286 8.4.7 – Structural Analysis .................................................................................................................288

8.5 – OTHER LOCOMOTION ......................................................................................................................289 8.5.1 – Sled Alternative ......................................................................................................................289 8.5.2 – Spring Launch Alternative .....................................................................................................290 8.5.3 – Ski Alternative ........................................................................................................................292

9 – ABOUT THE AUTHORS ................................................................................................................... 293

10 – REFERENCES .................................................................................................................................. 300

1 – Foreword

This report represents the culmination of an intensive spacecraft design course, A&AE

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) attitude control, communications, mission operations, 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 absolute cost of

delivering a small payload to the Moon's surface and to minimize the relative cost of

delivering an arbitrary payload. In both cases the payload was required to travel 500

meters across the lunar surface and to transmit images and data back to the Earth. In the

small payload case, the mission corresponds to the Google Lunar X PRIZE. In the case

of the large payload, the idea was to determine the payload mass that provides the lowest

cost per kilogram, i.e., “the biggest bang for the buck,” that would perhaps provide

insight into the size of a delivery system that could support a human outpost on the

Moon. While cost is very difficult to assess (in large part because of the proprietary

nature of the subject), the students managed to give meaningful and reasoned estimates of

the driving economic factors.

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.

Professor James M. Longuski

Purdue University

April 6, 2009

Introduction - Background Section 2.1,

Author: Solomon Westerman

2 – Introduction

2.1 – Background

The Moon has been a point of interest to humans since the beginning of recorded time. It

has been the subject of many works of art and literature and the inspiration for countless

others. While the Moon has always appeared tantalizingly close, advancements in science

and technology made only in the past fifty years have enabled lunar exploration.

Thrust forward by the space race between the United States and the Soviet Union in the

late 1950‟s and 1960‟s, robotic exploration of the Moon started off with a bang (or,

rather, a resounding “thud”). Both countries encountered recurring failures during their

early attempts at exploration, and they quickly understood the full complexity of lunar

missions. This sentiment resonated in President John F. Kennedy‟s speech at Rice

University on September 12, 1962, stating that “we choose to go to the Moon in this

decade and do the other things, not because they are easy, but because they are hard.”

Kennedy‟s ambitious goal of landing a man on the moon before 1970 established the

Apollo Program.

After the Apollo program achieved its primary objective in 1969, the program retired in

1972, initiating a long lull of worldwide lunar exploration. However, recent exploration

initiatives, emerging space programs, and potential lunar resources have placed the Moon

back in the “hot seat” of exploration.

To this day, lunar exploration has remained strictly the domain of large government-run

organizations due to prohibitive mission costs. The Google Lunar X PRIZE (GLXP)

aims to change this fact. A “$30 million international competition to safely land a robot

on the surface of the Moon, travel 500 meters over the lunar surface, and send images and

data back to the Earth,” the GLXP requires that “teams must be at least 90% privately

Introduction - Background Section 2.1,


funded” (Google). By providing a monetary prize, the X PRIZE Foundation aims to

enable commercial access to the Moon, bringing the Moon back to the forefront of

inspiration for individuals and corporations alike.

Introduction – What‟s in this Report? Section 2.2,


2.2 – What’s in this Report?

Project Xpedition is the product of Purdue University‟s Aeronautical and Astronautical

Engineering Department Senior Spacecraft Design class in the spring of 2009. This

report represents countless hours of work by 30 students over the course of a semester.

This report is divided into two sections: a Main Body and an Appendix. The Main Body

consists of a mission overview and important specifications for the team‟s vehicles. All

analysis and detail of the solution path are contained in the Appendix. The codes

referenced in this paper are available on the team‟s website.

Introduction - Acknowledgements Section 2.3,


2.3 – Acknowledgements

The design team would like to thank the following individuals for sharing their time,

expertise, and advice. Their continued support enabled us to create a quality design

fitting of the School of Aeronautics and Astronautics.

AAE450 Instructional Team

James M. Longuski, Professor – Instructor

Alfred Lynam, Graduate Student – Teaching Assistant

Joseph Gangestad, Graduate Student – Project Advisor

We would also like to recognize guest lecturers, industry contacts, and other professors

for their contributions to the project. Their guidance was invaluable in the design

process.

Guest Lecturers

David Filmer, Professor – Link Budget Analysis

Robert Manning, Graduate Student – Thermal Control

Dr. Boris Yendler, Lockheed Martin – Satellite Thermal Management

Industry and Academic Contacts

Edward Bushway III – Moog Inc.

Daniel Grebow – Graduate Student

Stephen Heister – Professor

Ivana Hrbud – Professor

Martin Ozimek – Graduate Student

Christopher Patterson – Graduate Student

Bruce Pote – Busek Company

Peter Van Beek – Maxon Motors

Introduction - Acknowledgements Section 2.3,


Additionally, we would like to thank LunaTrex, an official Google Lunar X PRIZE team

for the insight and guidance they shared with our team. We wish LunaTrex the best of

luck in the competition!

LunaTrex Lecturers

Pete Bitar – Team Leader

Mary Cafasso – Technical Team Leader

Joseph Gangestad – CIO

Introduction – Acronym List Section 2.4,


2.4 – Acronym List

This acronym list provides a comprehensive list of acronyms used in this paper in

alphabetical order. We hope this will prove a useful tool for you while reading this

report.

ACS – Attitude Control System

AOL – Active Oxygen Loss

CAD – Computer Aided Design

CCAFS – Cape Canaveral Air Force Station

CMOS – Complementary Metal Oxide Semiconductor

COTS – Commercial Off-The-Shelf

CPU – Central Processing Unit

CuInSe – Copper-indium-selenide

DC/DC – Direct Current/Direct Current

EP – Electric Propulsion

FCS – Flow Control System

FEA – Finite Element Analysis

FEM – Finite Element Model

FS – Factor of Safety

GaAs – Gallium-Arsenide

GaInP2 – Gallium-indium-phosphorus

Ge – Germanium

GLXP – Google Lunar X PRIZE

HD – High Definition

HET – Hall Effect Thruster

HETL – High Energy Tangent Landing

HRE – Hybrid Rocket Engine

LEO – Low Earth Orbit

Introduction – Acronym List Section 2.4,


LL – Lunar Lander

LLLE – Lunar Lander Locomotion Engine

LLME – Lunar Lander Main Engine

LLO – Low Lunar Orbit

MLI – Multi-Layer Insulation

MOSFET – Metal Oxide Semiconductor Field Effect Transistor

MS – Margin of Safety

NAND – Inverted AND logic

OTV – Orbital Transfer Vehicle

PAF – Payload Attach Faring

PARI – Pisgah Astronomical Research Institute

PCDU – Power Conditioning and Distribution Unit

PFA – Perfluoroalkoxyethylene

PPU – Power Processing Unit

PTFE – Polytetrafluoroethylene

RFHRE – Radial Flow Hybrid Rocket Engine

RTG – Radio-isotope Thermo-electric Generator

SNAP – System Nuclear Auxiliary Power

SRM – Solid Rocket Motor

TLI – Trans-lunar Injection (also Lunar Transfer)

TOF – Time of Flight

UHF – Ultra High Frequency

VIRAC – Ventspils International Radio Astronomy Centre

Project Overview – Design Requirements Section 3.1,


3 – Project Overview

3.1 – Design Requirements

The goal of Project Xpedition is threefold:

Project Xpedition Goals

1) Minimize the cost of a mission satisfying GLXP requirements – while

carrying a ballast of 100g.

2) Minimize the cost of a mission satisfying GLXP requirements – while

carrying a ballast of 10kg.

3) Minimize the cost per kilogram to the lunar surface satisfying slight

modifications to the GLXP requirements discussed below.

The first and second goals of Project Xpedition are similar to the goals of teams currently

participating in the GLXP. The ballast in either of these cases could be representative of

a small scientific or commercial payload, as commercial interest in a low cost payload

delivery service to the lunar surface is promising. Because the mission requirements for

the first and second goals are directly related to the GLXP, these two cases are referred

together as “GLXP-Sized” missions.

The third goal of Project Xpedition is distinctly different from the first and second goals.

The global minimum of cost per kilogram to the lunar surface satisfying slight

modifications to the GLXP rules is more applicable to a lunar base re-supply mission

than the GLXP contest itself. Because this design goal is open-ended on the range of

payload deliverable to the lunar surface, this goal is referred to as the “Large” mission.

We provide a summary of the GLXP requirements below. For a detailed listing of GLXP

requirements, please refer to Section A-9 in the Appendix.



Base GLXP Requirements

1) Safely land a robotic vehicle on the lunar surface

2) The vehicle or secondary vehicle deployed by the vehicle must move a

distance of 500 meters on the surface of the Moon in a deliberate manner.

3) Transmit an “Arrival Mooncast” and a “Mission Complete Mooncast” to

Earth.

The arrival mooncast and mission complete mooncast are specific sets of data to be sent

back to Earth in order to verify successful completion of the GLXP requirements.

Additional mission success requirements are imposed on each mission. These

requirements are provided by the instructor and are not part of the GLXP.

Mission Success Requirements

1) All three payload cases should have a 90% probability of success.

2) GLXP-Sized mission requirements must be satisfied by the GLXP deadline on

December 31, 2012.

We approach the probability of success requirement through different methods for the

GLXP-sized and Large payload missions. A cost-effective GLXP-Sized mission was

deemed impractical to satisfy the success rate with a single mission, so we plan multiple

launches to meet the overall success rate requirements.

Selected mission configurations for all three missions lead to flight times of

approximately one year. If a re-supply mission to a lunar base fails, a potential gap of

more than two years‟ resources is possible until the next mission can reach the base. For

this reason, the reliability requirement of the third goal is more stringent than the GLXP-

sized missions: a mission success rate of 90% is required for a single mission.



In addition to satisfying the base GLXP and mission success requirements, the instructor

placed further requirements on the mission designs:

Additional Requirements

1) The vehicle or any secondary vehicle must survive one lunar day and one

lunar night. After the lunar night, a vehicle must transmit eight minutes of

video to Earth. This optional GLXP bonus requirement is considered to carry

the full bonus monetary prize in addition to the main GLXP prize for the

GLXP sized missions.

2) The small “XPF Payload” described in the GLXP guidelines is ignored in

favor of ballast described in the Project Xpedition Goals.

3) A trajectory correction maneuver of 50 m/s provided by the main propulsion

system is required during trans-lunar injection.

4) The communications system during trans-lunar injection is required to be in

contact 90% of the transfer time.

5) The landing site must be near a historical landmark.

6) If using wheels to locomote on the lunar surface, the vehicle is required to

have the ability to climb a 45-degree incline, on a solid surface.

We make slight modifications to the base GLXP requirements and Additional

Requirements for the large payload case:

Large Payload Requirements

1) Any vehicle used to satisfy the large-payload requirements may “hover” at a

maximum of 100 meters above the lunar surface and translate 500 meters

without touching down first.

2) Any vehicle used to translate 500 meters must do so in over one minute.



We add these requirements with a lunar re-supply mission in mind. The first requirement

eliminates the requirement of touching down on the surface before locomotion, deemed

unnecessary for lunar re-supply. However, the 500-meter displacement requirement is

still in effect in order to allow for re-supply of a lunar base with small errors in landing

location by the re-supply vehicle. The second requirement reduces the risk of impact to

the lunar base by constraining the average approach speed of the re-supply vehicle.

Project Overview – Interpretation of Design Requirements Section 3.2,


3.2 – Interpretation of Design Requirements

Project Xpedition is a feasibility study aimed at providing an initial design concept as a

stepping stone to further development of low cost lunar missions. This limited scope

allows us to avoid obstacles encountered in detailed mission design. We acknowledge,

but ignore the following issues:

1) Politics

Our project involves physics, not politics. International treaties and export

control are ignored. In reality, red tape from a number of sources will cause a lag

in development that may push mission timelines outside acceptable limits.

2) Launch vehicle piggybacking

We assume any unused capacity in the launch vehicle can be sold back to the

launch vehicle company, as the Project Manager does not anticipate having

sufficient funds in his checking account to purchase the unused capacity. This is

a reasonable assumption, as launch vehicles typically carry multiple, smaller

secondary satellites in addition to a primary satellite. We further assume that our

selected launch vehicle will deliver our payload to a final delivery orbital altitude

of our choosing (within capability).

3) Launch timing

Launch vehicles are typically booked years in advance and have limited flexibility

in launch dates. In our analysis, we assume launch vehicles can launch at any

time from their primary facility.

Project Overview – Design Process Section 3.3,


3.3 – Design Process

We strive to maintain transparency in the design process in order to provide insight to our

selected mission configurations. Although we believe our selected mission configuration

is appropriate for our design process, a different design process is not guaranteed to

provide the same results.

The Project Manager and Assistant Project Manager are responsible for the successful

completion of the project. The Project Manager is not assigned a technical group but

instead acts as the primary systems engineer for the team. The Project Manager is

responsible for the structure of the design process as well as the final selection of the

mission configuration, total cost, and risk analysis. The Assistant Project Manager, in

addition to being assigned to a technical group, assists in the structure and management

of the design process.

Our design process was clearly divided into three steps: preliminary design, alternative

design and evaluation, and final design and evaluation. All three payload cases were

analyzed simultaneously.

Preliminary Design – Weeks 1 - 3

The team was initially divided into six disciplines: Attitude, Communications, Mission

Operations, Power, Propulsion, and Structures/Thermal. Each of these technical groups

roughly corresponds to specialty areas within the School of Aeronautics and

Astronautics. Each group member has special interest and extra coursework in their

group. Technical groups contain between four and seven members including a group

leader who was responsible for attending extra meetings and delegating work

appropriately within the group.



Design is inherently an iterative process that requires a seed to grow. The team began the

design with this seed – X0. This seed described a conventional lunar mission, based on

proven, historical methods to deliver payload to the lunar surface. While not optimized

and certainly not cost-effective, the X0 design provided a baseline for comparison to all

alternatives suggested later in the design process. This allowed each group to break

down systems on X0 and attempt to improve upon it, as detailed in the alternative design

and evaluation process.

Alternative Design and Evaluation – Weeks 3 - 7

In the third week, each team member was assigned a “phase group” in addition to a

technical group. Ignoring the Earth launch, three phases completely describe our lunar

mission: lunar transfer, lunar descent, and locomotion. Phase groups consist of one or

more members from each technical group and proved to be more effective and agile than

technical groups when evaluating alternative mission architectures. Phase groups avoid

some of the iterative nature of design: each phase group is relatively independent of each

other and usually only needed high level parameters from other phases to complete

analysis. As an additional benefit, phase groups were not locked in to a specific vehicle

type to complete a phase and can work with other phase groups to develop an efficient

architecture.

The primary action of the alternative design process was to develop and analyze

alternative methods within each phase. If an alternative was clearly superior to the X0

design, that alternative replaced the old design and became the new baseline. Each

alternative design was taken to a different level of maturity depending on its level of

merit within the mission as a whole.

At the end of the seventh week, the architectures for lunar transfer and lunar descent were

firm, with clear winners chosen from alternatives. Locomotion, however, went through a

rigorous process involving trades between mass, power, cost, and reliability to ensure the



lowest overall mission cost. The selection of the mission configuration for all phases at

the end of the seventh week allowed two weeks to finish final analysis on the selected

mission configuration for each payload case.

Final Design and Evaluation – Weeks 8 – 9

The purpose of final design and evaluation was to refine the models used in the

alternative design and evaluation process to provide reasonable accuracy to the selected

mission configuration. Changing mission configurations in this step was only done under

extenuating circumstances. The team formed “integration managers” to handle the fast

turnaround on numbers between phase groups during this process. Three integration

managers, each assigned to a payload case, managed the mass, power, volume, and cost

budgets for their payload case. By the end of the ninth week, finalization of the mission

configuration for each payload case was complete.

Project Overview – Risk and Cost Analysis Section 3.4,


3.4 – Risk and Cost Analysis

The concepts of risk and cost are interrelated and are thus combined in this section of the

paper. While true costs remain elusive and a detailed risk analysis is outside the scope of

this project, we believe the models used in this project are valid for a feasibility study.

The purpose of this section is to highlight the methodologies used in risk and cost

analysis in order to give further insight into our selected mission configurations.

Mission risk and cost are broken down by vehicle. Costs for each vehicle are broken into

the following categories:

1) Purchase – the cost of purchasing a component from another company.

2) Integration – the cost associated with manufacturing the vehicle.

3) Research and Development – the cost of developing or refining a technology in-

house.

Overhead is a category of cost that we include which is not associated with a particular

vehicle. Overhead represents the recurring cost of maintaining a company. This cost

includes salary for engineers working on the software and design levels, software license

costs, and communication rental expenses.

We assume that all purchases are made in 2009. Integration, research and development

costs are spread out evenly over three years and overhead over four years. We purchase

the launch vehicle in 2011, the same year as launch. All costs include a flat 3.8%

inflation rate used to represent all past and future cash flows in terms of 2009 dollars.

Our costing method tends to underestimate the total mission costs. Although an accurate

estimate of total mission cost requires more analysis, this does not affect the technical

feasibility of this project.



Risk analysis is also broken down by vehicle. The risk analysis technique used for the

Earth launch phase is distinctly different from the technique used for the other three

phases. Estimating the success rate of a launch vehicle is not trivial – the problem is

complicated further if the vehicle does not have extensive heritage. In our analysis, we

assume the launch vehicle probability of success is a function of historical reliability data

and the number of stages in the launch vehicle (a measure of complexity of the system.)

In the GLXP-sized payload cases we selected the Dnepr launch vehicle, a modified

version of an old intercontinental ballistic missile. The success rate of the launch vehicle

depends on the number of stages on the launch vehicle and historical launch data. For the

large-sized payload case, the selected launch vehicle has yet to be flown. The reliability

of this launch vehicle is assumed to be similar to the success rate of similar-sized launch

vehicles from the same country of origin.

Risk analysis for each vehicle that we design follows a different methodology from

launch vehicle risk analysis. We identify major failure modes for the vehicle and assign

an appropriate probability of success for each failure mode. Moving parts and systems

developed in-house are typically assigned a lower base reliability than non-moving or

purchased systems. If necessary, we augment this base reliability with money from

research and development to a maximum reliability cap. The probability of success for

each of the failure modes are multiplied together to calculate the success rate of each

vehicle. Each vehicle‟s probability of success is then multiplied together to produce the

overall mission success rate.

In some cases, the reliability of a particular system is too low to satisfy the mission

requirements, even with research and development. Our approach to satisfy the

requirements is to add a redundant system to the vehicle. This greatly increases the

success rate of the particular sub-system, as we assume the cause of a failure in one

system will not reduce the redundant system‟s success rate. This assumption leads to



optimistic success rates as the cause of failure of one system is potentially coupled with

the redundant system.

Project Overview – Results Section 4,


4 – Results

Our feasibility study found that it is practical for a commercial company (with a limited

funding base) to deliver small payloads to the lunar surface. However, a commercial

company is not likely to profit directly from the Google Lunar X PRIZE purse money.

Table 4-1 summarizes total mission costs (with the arbitrary payload case included as a

comparison to the GLXP-sized missions).

Table 4-1 Mission Results

Mission Payload Total Cost Cost/Payload Profit

GLXP 100 g $27.1M $271M/kg $(4.8M)

GLXP 10 kg $29.6M $3.0M/kg $(7.3M)

Large 1743 kg $222.6M $128k/kg ---

Footnotes: Payload – Payload mass deliverable to lunar surface

Total Cost – Mission cost, 2009 dollars

Cost/Payload – Cost per kilogram of payload

Profit – Net cash flow with GLXP purse

The most cost-effective mission configuration is relatively independent of payload

delivered to the surface. In order to minimize total mission cost, a company should use a

solar-electric propulsion system while the spacecraft travels from the Earth to the Moon.

In order to deliver payloads from lunar orbit to the lunar surface, a conventional, soft-

landing vehicle powered by a single hybrid engine is a cost-effective solution. Our

analysis indicates that this mission configuration results in low mission costs for payloads

ranging from 1 gram to approximately two metric tons. Although our analysis did not

converge on a minimum mission cost for payloads larger than two metric tons, the team

believes this mission configuration is an excellent springboard for further analysis of

large payload cases.

Project Overview – Results Section 4,


Our analysis indicates four feasible methods to travel short distances on the lunar surface

for a large range of payload sizes. Our cost model indicates the technique of locomotion

has a relatively low effect on mission cost. While we are confident in our selected

methods of locomotion, a different method of selection may lead to a different type of

locomotion. We split our locomotion configuration results into two categories: missions

satisfying the GLXP requirements and missions satisfying the modified GLXP

requirements as outlined in the Design Requirements section.

If a company is attempting to satisfy the modified GLXP requirements, the most cost-

effective design for translating across the surface to deliver payload to a lunar base is to

throttle the main landing engine to a low thrust level and translate above the lunar surface

after the main descent burn is completed. This method is an inexpensive and reliable

technique to displace short distances on the lunar surface. Our analysis indicates this

“hover” method is the global minimum in total mission cost for all payload cases.

However, this method is only incorporated in the arbitrary payload mission, as this

mission is not required to satisfy the exact GLXP requirements.

In order to satisfy the GLXP requirements and win the purse money, a different method

of locomotion is ideal and is dependent on the ballast mass prescribed in the mission

requirements. For a GLXP-sized mission with 100g ballast, the optimal method of

locomotion across the lunar surface is a small spherical ball with a movable, offset center

of mass. We use two identical spherical balls to maintain an acceptable level of

reliability. This type of locomotion is not ideal for the larger ballast case of 10kg, where

a separate chemical propulsion system becomes the preferred method of locomotion

across the lunar surface. We use a single hybrid engine with a separate combustion

chamber from the main landing engine in order to satisfy the GLXP locomotion

requirements. Like the 100g case, two independent systems of locomotion are required

to maintain an acceptable level of reliability.

Mission Configuration – 100g Payload Section 5,


5 – Mission Configuration 100g Payload

Mission Configuration – 100g Payload – System Overview Section 5.1,

Author: Brittany Waletzko

5.1 – System Overview

The mission configuration for the 100g payload consists of the orbital transfer vehicle

(OTV), Lunar Lander, and two redundant Space Balls that move across the lunar surface.

We design the system to move 500m across the lunar surface, survive the lunar night,

take and send pictures and video, and visit a heritage site on the moon. Figure 5.1-1

outlines the configuration of the entire system, and Fig. 5.1-2 shows the relative sizes

between the Space Balls and the Lunar Lander.

Fig. 5.1-1 Overview of 100g payload mission stack.

(Christine Troy)



Fig. 5.1-2 Space Balls and Lunar Lander on the Moon.

(Ryan Lehto)

Table 5.1-1 summarizes the mission timeline for the 100g payload. The landing site is in

Mare Cognitum, near the Apollo 12 landing site.



Table 5.1-1 Mission Timeline

Elapsed Time

(ddd:hh:mm) Event Vehicle

-365:00:00 Launch Launch Vehicle/OTV

000:00:00 Arrive in LLO OTV

000:00:03 In lower orbit Lander

000:00:04 Rotate and Land Lander

000:00:04 Systems check Space Ball

000:00:05 Deployment from Lander Space Ball

000:00:06 Orientation Space Ball

000:00:06 Travel 500m Space Ball

000:00:14 Braking maneuver, dust removal Space Ball

000:00:15 Take picture of Lander, Begin

transmission to Lander Space Ball

000:00:23 End photo transmission Space Ball

000:00:23

Transmit arrival Mooncast (near

real-time video, photos, HD video,

XPF set asides, data uplink set) to

Earth

Lander

001:33:56

Transmit Mission Complete

Mooncast (near real time video,

photos, HD video)

Lander

002:08:04 Finished transmitting, prepare for

night Lander

009:00:00 Standby for lunar night Lander

025:00:00 Power up after night Lander

026:00:00 Transmit telemetry and photo Lander

026:00:14 Mission Complete

Elapsed Time given in days, hours, and minutes

Table 5.1-2 summarizes system masses at key mission checkpoints.

Table 5.1-2 100g Payload System Masses

Parameter Mass Units

Injected Mass to Low Earth Orbit 436 kg

Injected Mass to Low Lunar Orbit 156 kg

Mass on Lunar Surface 79 kg

Payload Delivered 100 g

The following sections present greater detail on each vehicle subsystem.

Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.1,

Author: Zarinah Blockton

5.2 – Launch Vehicle

5.2.1 – Launch Vehicle Selection

We select the Dnepr-1 rocket to place the 100 g payload into a Low Earth Orbit (LEO).

The rocket is a converted intercontinental ballistic missile operated by the International

Space Company Kosmotras (a joint venture among Russia, Ukraine, and Kazakhstan).

The Dnepr-1 was chosen based on its low launch cost and adequate reliability. The

Dnepr-1 employs previous technologies and stages from its missile counterpart which

contributes to its low launch cost. To date, the Dnepr-1 has eleven successful missions

and only one failure. Figure 5.2.1-1 shows the launch profile of the Denpr-1.

Fig. 5.2.1-1 Dnepr-1 Launch Profile.

(Space Launch System Dnepr User’s Guide)



Table 5.2.1-1 Characteristics of the Dnepr-1 Launch Vehicle

Variable Value Units

Height

Diameter

Lift-off Mass

Mass to 400km orbit

Number of Stages

Orbit Inclination

Cost per kilogram to LEO

34.3

3

208,900

3,400

3

50.5

4,800

m

m

kg

kg

--

degrees

$/kg

The Dnepr-1 operates from one of the largest spaceports in the world, Baikonur

Cosmodrome in Kazakhstan. The Baikonur Cosmodrome is located a sufficient distance

from densely populated areas; this will ensure safety during stage drops. Although

launching from a site this far north of the equator decreases launch vehicle capability, the

launch price and performance in LEO still make the Dnepr-1 the most attractive option

for our mission configuration. Figure 5.2.1-2 shows the location of our launch site.

Fig. 5.2.1-2 Map of Launch Site.

(Google Maps)


Author: Andrew Damon

5.2.2 – Earth Parking Orbit Selection

Atmospheric Drag

We display the effects of atmospheric drag from altitudes of 200 km to 500 km in Table

5.2.2-1. At 200 km the drag force is still quite significant, and would cause the parking

orbit to decay quickly. Based on the data in Table 5.2.2-1, a parking orbit of 400 km is

the optimal drag solution. At 400 km the drag force is less than 5% of the thrust (thrust-

to-drag is greater than 20). The industry standard for minimum thrust-to-drag ratio in

Earth orbit is approximately 10. A higher orbit will only further increase our launch costs,

which are the largest mission cost driver. The 400 km altitude is also within the

capabilities of our selected launch vehicle for the 100g payload, the Dnepr.

Table 5.2.2-1 Atmospheric Drag Effects at Various Altitudes :

Altitude (km) Drag (mN) Thrust/Drag

200 76.4 1.44

300 17.8 6.18

400 4.10 26.83

500 0.96 114.6


Author: Levi Brown

Cost Comparison for Parking Orbit

The altitude of the parking orbit affects both launch and lunar transfer costs. Higher

parking orbits reduce the cost of lunar transfer by decreasing mass and power

requirements for the OTV; however, the launch cost increases because it is a higher

energy orbit. We find that launch cost has the greatest impact on overall mission cost, so

we select a parking orbit that minimizes launch cost. We see in Fig. 5.2.2-1 that

performing translunar injection from the lowest parking orbit possible results in

minimum cost.

Fig. 5.2.2-1 Mission Cost with Varying Parking Orbit Altitude.

(Levi Brown)

Because we must launch into a minimum 400 km orbit to overcome drag effects, we

select a 400 km circular parking orbit.

400 500 600 700 800 900 1000 1100 12004

6

8

10

12

14

16

18

Parking Orbit Altitude (km)

Launch a

nd L

unar

Tra

nsfe

r C

ost(

Mill

ion $

)


Author: Kristopher Ezra

5.2.3 – Attitude Determination in LEO

To maintain attitude control in low Earth orbit (LEO), we employ both a sun sensor and

star sensor. When the payload faring detaches from the launch vehicle, these sensors

come online and determine the attitude of the orbit transfer vehicle. The sun sensor is a

product of Valley Forge Composites and is accurate to approximately one degree. Sun

sensors are designed for attitude determination during de-spin maneuvers, such as the

OTV leaving the launch vehicle, as well as for directing solar arrays while in transit to

the Moon. The star sensor, used in parallel with the sun sensor, has much higher

accuracy and is used for attitude determination throughout the mission. This particular

sensor performs to an accuracy of 15 arcseconds (one fourth of a degree) by identifying

the location of stars instead of the sun. The combination of these subsystems can satisfy

our requirements for determining and monitoring attitude while the OTV remains in

LEO.

These systems have very high performance and use only a modest amount of spacecraft

resources. The sun sensor we selected has a mass of 0.35 kg and at peak power

consumption uses only 2.5 Watts. The star sensor is slightly more robust with a mass just

less than 3.2 kg, volume of 16,400 cubic centimeters, and a maximum power

consumption of 10.2 Watts. These units together have a total mass of approximately 3.55

kg and determine the attitude for the entire mission (Valley Forge, 2009). Furthermore,

the sensors can be purchased from Valley Forge Composite Technologies in a package

including reaction wheels for approximately $400,000.

We must maintain continuous contact with ground stations after the OTV deploys; an

accuracy in attitude determination of approximately one degree ensures we can

accomplish this goal. This constraint that a ground station must be in view at all times

would place a hefty burden on attitude determination and attitude control except that we

employ a moveable antenna. For this reason, we assume that the upper bound on

required accuracy (the higher the accuracy, the lower the value in degrees) is about half a



degree. This accuracy is definitely achievable because the star sensor alone has an

accuracy of one fourth of a degree. In typical cases, it is reasonable to think that in LEO

even the sun sensor could be accurate enough to place a ground station in view of the

antenna.

Unlike other subsystems onboard the OTV, the attitude determination systems do not

scale or vary with changing mission requirements. As we will discuss later in this

document (Sections 6.2.3 and 7.2.3), there is no change to the attitude determination

methods regardless of new payload masses, new landing/locomotion equipment, or new

mission architectures. The star sensor and the sun sensor will be employed without

significant changes (excluding perhaps position within the OTV) in each of the three

payload cases.

Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1,

Author: Tim Rebold

5.3- Lunar Transfer

5.3.1 – Structure

Limit Loads

Launch loads or limit loads from different phases of the launch are obtained from the

Dnepr User‟s Manual. We choose these loads to evaluate the capability of our spacecraft

design. Table 5.3.1-1 shows these launch loads for the selected launch vehicle. Loads

are expressed in terms of gravitational acceleration (g‟s), and produce forces in all

structural elements of our spacecraft. These forces are broken into axial and lateral

components throughout launch. We select the highest levels of acceleration to apply to

our spacecraft design (axial and lateral simultaneously). We base our entire structural

analysis on these selected loads. This assumption is valid because the most severe

dynamic and static loads are expected to occur through the launch phase of the mission

for our Orbital Transfer Vehicle (OTV).

Table 5.3.1-1 Dnepr launch limit loads

Event

Axial

Acceleration

Lateral

Acceleration

1st Stage Burn: Maximum Lateral Acceleration 3.0 ± 0.5 0.5 ± 0.5

2nd Stage Burn: Maximum Longitudinal Acceleration 7.8 ± 0.5 0.2

Table based from Dnepr User‟s Guide

The loads specified are valid only if the spacecraft meets certain stiffness, or natural

frequency requirements. These requirements are summarized in Table 5.3.1-2 for the

Dnepr launch vehicle. These requirements are based on the launch vehicle‟s dynamic

characteristics and are used to prevent an amplification of the loading.

Table 5.3.1-2 Dnepr spacecraft stiffness requirements

Thrust (Hz) Lateral (Hz)

20 10

Table based from Denpr User‟s Guide


Author: Tim Rebold

Margin of Safety

We apply a factor of safety (FS) when we conduct our analysis. This is done to ensure

that uncertainty will not make our analyses invalid. Uncertainty can come in the form of

imperfections in the structure, material property variations (strength), manufacturing

tolerance errors in structural members, errors in predicting loads in the mission, and other

factors. The yield margin of safety (MS) which we use in all of our analyses is shown as:

𝑀𝑆 = 𝐴𝑙𝑙𝑜𝑤𝑎𝑏𝑙𝑒 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠)

𝐷𝑒𝑠𝑖𝑔𝑛 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠) − 1 (5.3.1-1)

The design yield load takes the predicted loads or stresses and multiplies it by the FS

chosen for the mission. If the margin of safety is negative it means that we have

exceeded the allowable loading or stress of the material or member. This means we must

redesign or find a way to lower the expected loads. We use a FS of 1.5 for our

preliminary analyses to be conservative.

Configuration

Our OTV is 1.29 m tall including the Lunar Lander (LL) integration skirt, and has a 1.8

m diameter. Figure 5.3.1-1 displays the primary components and dimensions of the OTV

in a schematic. The OTV is cylindrical with its payload (Lunar Lander-cone shaped)

resting on top. An integration skirt joins the two. We select the overall dimensions of the

OTV based on sizing constraints. Once we choose the basic layout, we can select the

primary structural components that will make our spacecraft. In the following sections

we describe the primary structural components of our spacecraft.

Stiffeners

We use stiffeners as the primary support structure for the spacecraft. The stiffeners are

also referred to as stringers or C-Channels (due to their cross section shape). We choose

stiffeners to provide the main load path for all the forces seen by the spacecraft

throughout launch. The stiffeners are capable of withstanding the bending and


Author: Tim Rebold

compressive forces transferred throughout the entire structure. Six stiffeners provide

enough stiffness and strength as well as distribute the loads uniformly into the launch

vehicle interface at the aft end of the spacecraft

Fig. 5.3.1-1 OTV configuration and design showing primary components and dimensions.

(Tim Rebold)

Propulsion Module

The propulsion system of the OTV is supported with a truss frame structure. The

propulsion system consists of the Xenon propellant tank, feed system, Hall thruster, and

thermal protection equipment. Section A-5.3.1 provides more details on how we chose a

layout and sized the frame. Four members are the fewest needed to support our

propulsion system.

Electronics Module

The electronics module (E-MOD) houses all the electronics needed for the mission. The

structural support for the E-MOD consists of a floor support beam design and skin which


Author: Tim Rebold

covers the entire assembly. The skin serves as a barrier between the Sun‟s radiation and

equipment. The skin also offers micrometeorite protection during Lunar Transfer. We

also choose the skin overlay on the floor to act as a radiator to conduct unwanted heat

away from the electronics. The amount of space needed to fit all of the equipment

determines the allowable dimensions. Since the floor skin overlay is so thin (0.5 mm), it

is not expected that they will carry any loads. All equipment is mounted above the six

beam supports. The geometry and layout of this design is shown in Section A-5.3.1.

Skin

We place 0.5 mm skin (shear paneling) around the entire OTV to mainly provide

shielding from the Sun‟s radiation and to offer micrometeorite protection. We choose the

thickness to be as thin as possible to reduce weight yet provide adequate support for the

above reasons mentioned. From a structural standpoint, the skin will offer substantial

torsional rigidity for the entire OTV, and handle most of the shear loading seen in the

OTV.

Integration

Integration structure includes the Payload Attach Fitting (PAF) and LL skirt. The latter

joins the OTV to the LL, while the former joins the OTV to the launch vehicle inside the

fairing. There are six locations on both the base and top of the LL skirt where the OTV

and LL are bolted into. These six locations correspond to the location of each stiffener.

Similarly, the PAF attaches the OTV to the launch vehicle inside the fairing. The Dnepr

User‟s Guide shows a standard launch vehicle interface. This interface determines the

base dimensions of our PAF. Geometry can be seen in Section A-5.3.1.

Sizing

We size structural members in the OTV by predicting the expected loads transferred

through the spacecraft. We reduce the weight of the spacecraft by minimizing the cross

section dimensions of each member in the OTV. Modes of failure need to be well


Author: Tim Rebold

understood to identify the most likely cause for failure in a given member. The most

likely cause for failure will determine the criteria for our design. We can apply basic

techniques and equations to optimize the capability of each member, while making it as

light as possible. The most common failure modes that are used to determine the size of

structural members in our OTV design are discussed in detail in Section A-5.3.1.


Author: Korey Lemond

Material Selection

The OTV is generally comprised of three different materials. All structural frame

members are fabricated from aluminum due to the fact that they must be designed to

survive the launch loads that are applied to the vehicle. Aluminum has a higher yield

strength, and all structural frame members are therefore stronger. The shear paneling that

encompasses the vehicle was fabricated out of a lower strength magnesium alloyed with

small concentrations of aluminum and Zinc. The yield strength of the magnesium was

approximately two thirds of that of aluminum, but the density was also two thirds the

density of aluminum. Since shear panels in general carry little in the way of loads, instead

acting as a pathway for shear flow, we chose a lower strength lower density material in

the form of magnesium. Magnesium also is used as the material for the floor of the

electronics module and the floor of the OTV itself. The floor panel of the OTV carries

little in the way of loads due to the fact that the propulsion system is attached to the truss

frames. However, some form of meteorite protection is necessary, and magnesium

suffices. The third material present in the OTV is that of AIS 1015 low carbon steel. This

material is used in all pins, fasteners, and other connection devices, as these devices are

the only load path between structures and therefore see all the launch loads at a given

time.

Table 5.3.1-3 Material Properties

Material Density

(kg

/m3)

Yield

Strength (Pa)

Young‟s

Modulus (Pa)

Al 6065 2700 1.03E8 6.8E10

Mg AZ31 1770 1.5E8 4.4E10

AIS 1015 Low

Carbon Steel

7800 2.55E8 2.05E11


Author: Tim Rebold

FEA Analysis

As the configuration and sizing of the Orbital Transfer Vehicle (OTV) reaches maturity

we perform a detailed structural analysis of our model.

Fig. 5.3.1-2 Finite element model (FEM) of the OTV.

(Tim Rebold)

The software we use to perform our finite element analysis is IDEAS 12. We use this

package to perform a basic static load analysis to determine critical stresses and

displacements in structural members. We also perform a modal analysis to determine if

our spacecraft meets the stiffness requirements provided in the Dnepr User‟s Guide.

Figure 5.3.1-2 shows a complete finite element model (FEM) of our OTV spacecraft

Lander Skirt Analysis

The first analysis we perform is on the Lunar Lander (LL) skirt. By performing a basic

static test we can see the capability of the skirt. Figure 5.3.1-3 shows the LL skirt

geometry.


Author: Tim Rebold

Fig. 5.3.1-3 Lander skirt geometry.

(Tim Rebold)

Results

After performing the analysis we find that stress is not a concern. By picking the

locations of where the LL was integrated to the OTV we effectively limited stress and

displacement in the skirt. We find that a buckling analysis determines how thin we can

make our skirt walls (or webs). We settle on a final wall thickness by assuring ourselves

that the skirt does not buckle under the predicted loading.

OTV Analysis

We analyze the entire OTV next. We perform a system by system analysis to simplify

our results. We conduct a basic static test, and modal analysis to find the natural

frequencies of our spacecraft while mounted to the launch vehicle interface.

Results

The first system we observe is the propulsion module frame. The max Von Mises stress

is 324 N/mm2 and max displacement is 1.36 cm. This corresponds to a margin of safety

of -0.17. The negative margin of safety means that yielding occurs in that area of peak

stress. Material yield is unacceptable by our mission requirements. The member failing

is not seeing unreasonably high stresses, so a slight change in design should eliminate the

problem. Figure 5.3.1-4 shows the results for the propulsion frame. The margins of

safety for all other systems are passing.


Author: Tim Rebold

Fig. 5.3.1-4 Propulsion frame stress contour results indicate material failure.

(Tim Rebold)

We next perform a linear buckling analysis. The buckling load factor from this analysis

is 0.19, which means that a buckling failure occurs at the current level of applied loads.

We observe that a stiffener is first to buckle from the applied loads. We need to redesign

so this does not occur. We last perform a modal analysis. We observe that the first

lateral mode is well below the requirement. This means that the OTV needs to be stiffer

in the lateral direction.

Design Modifications

Next, we make modifications to our spacecraft until all requirements are satisfied.

OTV Modified Analysis Results

Again, the first system we observe is the propulsion module frame. The max Von Mises

stress is 80 N/mm2 and max displacement is 2.14 mm. This corresponds to a margin of


Author: Tim Rebold

safety of 2.14. The positive margin of safety means that this system meets mission

requirements. A buckling load factor of 1.42 and lateral mode of 10.6 Hz are also

determined from the buckling and modal analyses. Therefore, our OTV design meets all

structural requirements. Table 5.3.1-4 shows the structural component masses as a result

of these analyses.

Table 5.3.1-4 Component mass totals for OTV design

Components Mass (kg)

PAF* 47.04

E-MOD floor beams & overlay 5.25

Shear / Skin Panels 15

Propulsion Support Frame 3.01

Stringers / Stiffeners 12.48

Lander Skirt 12.19

Fasteners (welds, rivets, bolts, adhesives) ** 2.01

TOTAL 49.94

* Not included in final OTV mass

** Estimate


Author: Levi Brown

5.3.2 – Mission Operations

Trajectory

Per the analysis described in Section A-5.3.2, we design the trajectory for the 100 g

payload to have the configuration detailed in Table 5.3.2-1.

Table 5.3.2-1 100 g Payload Trajectory Configuration

Parameter Value Unit

Payload Mass 156.2 kg

Thrust 78.5 mN

Mass Flow Rate 4.1 mg/s

Earth Phase Angle 108 deg

Moon Phase Angle 288 deg

Parking Orbit Altitude 400 kg

Capture Orbit Altitude 25 kg

Initial Mass 436.0 kg

Flight Time 365 days

This model inherently contains a mismatch in position and velocity at the intersection of

the spiral out and spiral in curves (See Fig. 5.3.2-1). We calculate the propellant mass

required to produce the ΔV mismatch operating the main engine. We add this propellant

mass to the OTV to account for the error in position and velocity. We assume that

performing small maneuvers throughout the trajectory eliminates the fairly large bias in

position and velocity at the intersection point. Table 5.3.2-2 contains these mismatch

values.

Table 5.3.2-2 100 g Payload Trajectory Configuration

Parameter Mismatch Unit

Position 5676 kg

Velocity 417.6 km/s

Propellant Mass 9.4 kg

Figure 5.3.2-1 illustrates the resultant trajectory.


Author: Levi Brown

Fig. 5.3.2-1 100g Payload Trajectory.

(Levi Brown)

-2 -1 0 1 2 3 4

x 105

-2.5

-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

x 105

xhat

(km)

yhat(k

m)

Spiral Out and In of Spacecraft

Spiral Out from Earth

Spiral In to Moon


Author: Kara Akgulian

Lunar Capture Orbit

The lunar capture altitude is 25 km, which optimizes mass savings while maintaining

orbital stability. A small altitude produces the most efficient lunar descent trajectory and

propellant mass savings. However, the lunar terrain and the instability of these smaller

orbits restrict the altitude to the minimum value of 25 km. Once we enter this orbit the

Lunar Lander separates from the OTV and begins its descent to the Moon‟s surface.


Author: Levi Brown

Trajectory Correction Maneuver

One requirement for this project includes the capability to perform a 50 m/s burn for

course correction. Due to factors beyond the scope of this project, such as Sun

perturbations, the spacecraft will deviate from our trajectory design during actual flight.

To account for this bias, we carry additional propellant for the main engine. The

propellant is available for making small corrections in flight as necessary.

We calculate the propellant necessary for a 50 m/s burn and record the results in Table

5.3.2-3.

Table 5.3.2-3 Correction Maneuver Configuration

Parameter Value Unit

Isp 1952 sec

mo 436.0 kg

Propellant for Correction 1.1 kg


Author: Brad Appel

5.3.3 – Propulsion on the Orbital Transfer Vehicle

Propulsion Subsystem Overview

We select a low-power Hall Effect Thruster (HET) for primary thrust onboard the Orbital

Transfer Vehicle. HETs are a type of electrostatic propulsion (EP), and there are two

defining features to keep in mind with these devices: They have very low thrust (tens of

milliNewtons), and very high specific impulse (thousands of seconds). Our HET

propulsion system contains four critical components, including the Power Processing

Unit (PPU), the propellant tank, the propellant Flow Control System (FCS), and the

thruster itself. We place each component within the spacecraft as shown in Fig. 5.3.3-1.

Fig. 5.3.3-1 Placement of the key propulsion system components in the overall vehicle.

(Brad Appel)

The EP system propels the spacecraft from Low Earth Orbit to Lower Lunar Orbit with a

slow, continuous thrust. Our payload is the wet mass of the Lunar Lander in addition to

the Space Balls. Table 5.3.3-1 lists the required EP system totals for the 100g mission.

Table 5.3.3-1 Electric Propulsion System Totals


Wet Mass 169 kg

Dry Mass 30 kg

Required Power 1,540 Watts

Burn time 365 days

Purchase Cost $1.03 million 2009 dollars

2 m


Author: Brad Appel

The Hall Thruster

A Hall Effect Thruster provides three fundamental advantages for our mission. First, the

system has a very high payload mass fraction: 38% for our 100g mission. Second, HETs

physically require a small propulsion system - the thruster itself is no bigger than a

computer monitor, and its associated power conditioning equipment is about the size of a

PC tower. This compact feature drives down the volume as well as the mass of the

spacecraft. And third, the Hall Thruster system is inherently simpler than many other

propulsion options. This fact goes a long way in making the system cheaper to integrate

and more reliable overall.

Specifically, we use the BHT-1500 from the Busek Company for the main engine. There

are actually only a few commercial sources for Hall Thrusters, and from what data are

available, the BHT-1500 offers the best performance for our mission. Table 5.3.3-2

summarizes the performance and physical specifications of the thruster (Azziz , 2005).

The performance numbers (mass flow rate and specific impulse) are optimized by the

team in order to provide the cheapest overall system. For the best combination of high

power-to-thrust ratio and easy handling properties, we select Xenon as the propellant.

Table 5.3.3-2 Specifications for the Hall Thruster


Thrust 78.5 mN

Specific Impulse 1950 s


Power Input 1526 W

Efficiency 0.53 --

Input Voltage 350 VDC

Mass 5.7 kg

We place the HET protruding through the OTV floor paneling at the aft end of the

spacecraft. The thruster produces 700 Watts of excess heat, which we radiate away using

this Magnesium floor panel.


Author: Brad Appel

Power Processing Unit

The Hall Thruster comes with very specific and demanding power requirements. We

require a Power Processing Unit (PPU) to transform, monitor, and condition the power

input to the HET. Because the development of a new flight HET is usually accompanied

by the customization of a PPU, we logically select the Power Processing Unit from

Busek. Some important features of the PPU are listed in Table 5.3.3-3 (Osuga, 2007),

(Kay, 2001).

Table 5.3.3-3 Specifications for the Power Processing Unit


Operating Voltage 28 VDC

Input Power 1352 Watts

Efficiency 93% --

Mass 10 kg

The PPU serves other system functionalities as well. It collects all telemetry from the

HET, such as temperature and voltage data, and communicates with the spacecraft main

computer via a serial transmission cable. The PPU is also equipped to monitor and

command a mass flow controller, should the mission require it. The PPU has three critical

interfaces: Main power from PDCU to PPU, Commands and Telemetry from the central

computer to the PPU, and power input to the HET. These are all soft harness connections

and do not require a particular positioning for the PPU.

Within the spacecraft, we place the PPU forward of the rest of the propulsion system,

inside the OTV electronics compartment. This placement provides the most convenience

for thermal control and alignment of the spacecraft center of mass.

Xenon Propellant Storage and Flow Control System

We store the Xenon propellant in a supercritical state in order to minimize the volume it

takes up (supercritical means the gas is compressed to the brink of becoming a liquid).

This storage condition has two consequences: the tank must be strong enough to

withstand high pressure, and the tank must be temperature regulated to keep the Xenon


Author: Brad Appel

supercritical. We use a metal-lined, carbon composite overwrapped tank from Arde Inc.

The tank is spherical with a diameter of 60 cm and a dry mass of 12 kg. We initially

pressurize the tank to 15.2 MPa (2200 psia); however the pressure steadily drops

throughout the mission. For thermal control, a 5-Watt resistance heating wire is

embedded in the tank with thermostat regulation. We wrap the tank with Multi-Layer

Insulation in order to mitigate heat loss by radiation.

The tank feeds our Flow Control System (FCS), which conditions the Xenon flow to the

mass flow rate required by the Hall Thruster. Our FCS accomplishes four essential tasks:

First, a solenoid valve isolates the high pressure Xenon in the storage tank. Upon this

valve opening, the Xenon flows to a pressure regulator, where it is stepped down from

several thousand psi to around 40 psi (Moog Corporation, 2009). After passing through a

high-purity filter, the propellant flow is split into the three required feed lines for the

HET. The last stop for the Xenon before entering the thruster is through a flow restrictor,

which we calibrate to deliver the desired mass flow rate (Mott Corporation, 2009). These

flow restrictors are sensitive to temperature changes, and so we setup a heating coil and

thermostat system for the FCS as well.

Figure 5.3.3-2 is a schematic of the propulsion system on the OTV. A parts list is

included in Table 5.3.3-4.


Author: Brad Appel

Fig. 5.3.3-2 Electric Propulsion System Schematic.

(Brad Appel)

Table 5.3.3-4 Parts List for the EP System.

Label # Component

1 Tank Heating Wire

2 Xenon Storage Tank

3 Tank Multi-Layer Insulation

4 Solenoid Latch Valve

5 Pressure Regulator

6 High Purity Filter

7 Sintered Flow Restrictors

8 FCS Heating Wire

9 Thruster Radiation Panel

10 Hall-Effect Thruster

11 Power Processing Unit


Author: Brian Erson

5.3.4 – Attitude

Attitude Control

Spacecraft attitude control is necessary to maintain a specific trajectory within the

translunar phase of our mission. We are controlling the Orbital Transfer Vehicle (OTV)

by a set of three systems: sensors, reaction wheels, and thrusters. These systems comprise

our attitude control system (ACS) and combine to provide the necessary attitude

corrections to accomplish our mission.

Our inertial and relative positions must be known before any attitude correction can be

done. We use a set of sensors to obtain and translate this knowledge. The star sensor,

made by Valley Forge Composite Technologies (VFCT), takes pictures of visible stars

and compares them to a database of star systems. Since stars are considered inertially

stable, the orientation of the OTV relative to space can be determined.

A Sun sensor, also made by VFCT uses pictures of the Sun to obtain the position of the

OTV relative to the Sun; therefore, determining the position of the spacecraft relative to

the Earth and Moon. As sensing is independent to the size of a spacecraft, we use these

same sensor systems for all missions. These systems, however, only provide a picture of

the OTV in space, and provide no control to the spacecraft.

Reaction wheels comprise one layer of attitude control within the OTV. A reaction

wheel uses electric motors that spin weighted plates to create torque. This torque is used

to make small corrections in the attitude of the OTV. Our friends at VFCT also

manufacture reaction wheels. The VF MR 4.0 has enough capacity to overcome all

torques caused by environmental perturbations, thrust misalignment, and drift error. The

torque experienced has a cumulative effect and eventually will need to be de-spun to

recharge the torque-creating power of the reaction wheel.


Author: Brian Erson

Comprising the final layer of control, thrusters provide the force necessary to de-spin

reaction wheels and make independent attitude corrections. Several systems were

analyzed according to their cost, inert mass, and propellant mass required. We choose a

hydrogen peroxide (𝐻2𝑂2) monopropellant thruster to provide the counter torque needed

to de-spin the reaction wheels. The small thrusters are configured in a 4-axis redundant

system to provide reliable 3-axis control.

Our three attitude subsystems: sensors, reaction wheels, and thrusters combine to create a

complete system accomplishing the mission of controlling the OTV during the translunar

phase. Total attitude system mass and cost is small compare to mission entirety. VFCT is

providing a cost decrease pending the use of advertisement within our mission. Our OTV

provides more than enough power for mission success. System volume is small enough to

be easily integrated into the current OTV platform. Table 5.3.4-1 provides an overview

of our ACS specifications for the 100g payload.

Table 5.3.4-1 OTV ACS Budget for 100g payload

Device Mass (kg) Cost ($) Power Required(W)

VF STC 1 (star sensor) 6.4 133,333 20.4

VF SNS (sun sensor) 0.7 133,333 5

VF MR 19.6 10.4 133,333 76

𝐻2𝑂2 thruster 0.36 1,500 --

𝐻2𝑂2 propellant 0.02 100 --

Inert Mass 1.9 1,000 --

Totals 19.8 403,000 101.4



Attitude Propellant Use

Four factors are used in determining the type of propellant within the ACS: mass

implications, cost of total system, current use of propellant on OTV, and ease of

integration.

As seen in Table 5.3.4-1, we choose 𝐻2𝑂2 for its low weight and cost. The integration

requires a small increase in inert mass, but not enough to eliminate 𝐻2𝑂2 from

contention. The propellant is not currently used on the OTV, but the preceding factors

outweigh the possibility of using a different propellant. 𝐻2𝑂2 is a stable, nontoxic

substance that has long term storability. The performance characteristics outlined in

Section A-5.3.4 are adequate to provide enough thrust to fulfill all mission requirements.

Space Environment Perturbations

In this section we perform a “worst case” needs assessment for attitude control due to

environmental perturbing forces during the Lunar Transfer phase. In order to compute the

maximum expected torques, a few assumptions are made. The largest assumption in this

analysis is that all forces add linearly and are of their greatest magnitude. Additionally,

both a right circular cylinder and a cube are used to approximate the shape of the OTV so

as to calculate environmental forces on it. Many variables, such as the spacecraft's

electric charge and reflective properties, are historically based on missions including

Spirit-1 and Lunar Surveyor (Berge, 2009). The sources of environmental disturbances

include electromagnetic radiation from the sun, reflected radiation from the Earth, Earth's

thermal radiation, an estimated amount of spacecraft radiation, solar wind, gravity

torques, meteoroids, Earth's magnetic field, and a nonpropulsive mass expulsion from the

spacecraft (Longuski, 1982)(deWeck, 2001). We then compute the corresponding torques

based on the moment arm for each force on the OTV (see environmental.m).

Even with these conservative estimates, the sum of all the torques is only on the order of

1.5x10-3

N-m for the cylinder and 1.5x10-5

N-m for the cube. If we constantly use an



attitude control thruster to correct this torque, the cube only requires roughly 0.3 kg of

propellant (see environmentalpropmass.m). This compares to a historical average of

roughly twelve kilograms of attitude control propellant for spacecraft (deWeck, 2001).

As such, these torque corrections are small enough that they can be compensated for with

a reaction wheel.


Author: Joshua Elmshaeuser

5.3.5 – Communication

Communication Hardware/Configuration

We find that during the Lunar Transfer it is necessary to maintain communication

between Earth and the OTV for 90% of the time. To make this happen we must consider

the Lunar Transfer trajectory and the orientation of the vehicle to insure the antenna is

pointed in a direction capable of sending and receiving a signal. We conclude that to meet

these requirements one system is all that we need. Instead of placing a system on the

OTV, which would be abandoned during Lunar Decent, we place one system on the

Lander. This one system fulfills all communication requirements for Lunar Transfer,

Lunar Decent, and for operations while on the surface of the Moon. More information

regarding the selected system and configuration can be found in the lunar decent Section

5.4.5.


Author: Trenten Muller

Communication Ground Stations

In order to ensure our spacecraft behaves as intended we need to monitor it during the

Lunar transfer. In order to monitor our spacecraft we will make use of ground stations on

Earth. These ground stations and our spacecraft will communicate with each other using

radio antennae operating at 2.2 GHz in the S-Band range. Figure 5.3.5-1 which shows the

location of our ground stations as if the Earth is viewed looking at the North Pole as well

as the coverage “net” created.

Fig. 5.3.5-1 Our four ground stations. 100% tracking capability outside straight lines.

(Trenten Muller)

We use four ground stations. They are:

1) Mt. Pleasant Radio Observatory. Hobart, Tasmania, Australia. A 26 meter dish.

2) Hartebeesthoek Radio Astronomy Observatory (HRAO). Johannesburg, South

Africa. A 26 meter dish.

3) Pisgah Astronomical Research Institute (PARI). Rosman, North Carolina. USA.

One of the 26 meter dishes.

4) James Clark Maxwell Telescope. Mauna Kea Observatory, Hawaii, USA. A 15

meter dish.



The latitude and longitude for these four ground stations are listed in Table 5.3.5-1.

Table 5.3.5-1 Geographic Locations of Ground Stations

Ground Station Latitude (o) Longitude (

o)

Mt. Pleasant 42.81 S 147.44 E

HRAO 25.55 S 27.68 E

PARI 35.20 N 82.87 W

Maxwell 19.82 N 155.48 W

The four ground stations that we select will be used for all payload cases and will be used

during all mission phases. All have horizon to horizon scanning, meaning that they can

scan the entire sky at their locations, excluding geographical constraints. The James Clark

Maxwell Telescope will only be used to bridge the large gap between Mt. Pleasant and

PARI while within 8,400 km. The tracking ability of these stations will be above 90%.

Based on an e-mail conversation with Professor John M. Dickey (2009) of The

University of Tasmania, home to the Mt. Pleasant Observatory, we have estimated the

cost to operate these stations at $1 million for one year of usage. Professor Dickey

provided a quote of $1 million for the use of the Mt. Pleasant Observatory for one year of

continuous use. Because we will only use one station at a time we will be able to spread

the $1 million quote amongst all of the stations as we use them individually throughout

the year.


Author: Michael Christopher

Communications Link Budget

We budget 0.5 kbps (kilobytes per second) for uplink and downlink communication with

the Orbital Transfer Vehicle (OTV) during the long trip to lunar orbit. Our

communication link consists of telemetry transmission from the OTV, as well as

commands from Earth to the OTV. The telemetry from our OTV consists of vital data

concerning the spacecraft‟s operation, including information about attitude, power

consumption, and temperature. The commands that our OTV receives from Earth contain

data pertaining to mission operations procedures.

The 0.5 kbps we allocate for the communication link during the Lunar Transfer phase of

the mission is well below the bandwidth capabilities of our chosen communication

equipment. We choose not to employ this equipment to its full capability in an effort to

reduce power consumption during this lengthy voyage to lunar orbit. This electricity

conservation is vital to our mission, because of the power consumption of our electric

propulsion system and the premium of available power.

Communications Antenna Pivot

To communicate with ground stations during the Lunar Transfer phase of the mission our

spacecraft employs two patch antennae which are located on the Lunar Lander. We

choose to use the communication equipment on the Lunar Lander in order to eliminate

the need for redundant systems on the OTV. The elimination of unnecessary redundant

systems provides for a large savings in mass and cost.

As our spacecraft travels to lunar orbit its attitude in space will not always be conducive

to transmitting and receiving data with Earth ground stations. The high gain patch

antennae located on the Lunar Lander has a limited beam width, and needs to be pointed

towards Earth. As maneuvering the entire spacecraft for communication would be very

costly to the mission, we choose to simply gimbal the antennae. To do this gimbal



activity we have designed a patch antenna pivot. This pivot design is pictured in Fig.

5.3.5-2.

Fig. 5.3.5-2 Patch antenna pivot.

(Michael Christopher)

As see in Fig. 5.3.5-2 the pivot consists of two plates, a base plate and an antenna mount

plate, two connecting arms, and two stepper motors. The base plate is mounted directly

onto the outer surface of our Lunar Lander. We choose to include two precession stepper

motors to first increase range of motion, as seen in Fig. 5.3.5-3 and Fig. 5.3.5-4. Second,

we choose to include two motors for redundancy. If a single motor fails, the other motor

will still be able to operate the pivot with a slightly limited range of motion.



Fig. 5.3.5-3 Pivot in "normal" position.


Fig. 5.3.5-4 Pivot in 180 degrees from "normal" position.


The pivot integrated onto our Lunar Lander is displayed in Fig. 5.3.5-5; there is a

duplicate antenna/pivot assembly on the other side of the Lander. We include two of

these assemblies for yet more needed redundancy. With all of this redundancy we can

operate a communication link with three of the four motors out of operation, and one of

the two antennae out of operation. Without this redundancy, the communication system

would be a single point mission failure.



Our choice not to include a communication system on the OTV, and use the one on the

Lunar Lander, presented a reliability challenge. This challenge was overcome by

including redundancy both in the pivot design, and in the placement of two identical

systems. At the extremely low cost of approximately $83.50 and low mass of about 0.2

kg, including two systems is wise design choice.

Fig. 5.3.5-5 Pivot integrated onto the Lunar Lander.

(Josh Elmshaeuser)


Author: Ian Meginnis

5.3.6 – Thermal Control

Electronics Thermal Control

Three main components of the OTV need to be thermally controlled. These consist of the

electronics, the xenon tank, and the electric thruster. Because of the different thermal

requirements of each of these systems, each has a separate thermal control system. The

electronics comprise of the PCDU, the battery, the PPU, and the DC/DC converters.

Together, these components put out a total of 217W of heat. To cool these components,

we mount them on an aluminum plate, which removes the heat through conduction. We

note that the plate also serves as a structural support for the components. Attached under

the plate is a series of capillary heat pipes, with ammonia as the working fluid. When the

ammonia heats up and boils, the ammonia vapor travels to two radiators that are exposed

to space. After the heat radiates to space, the ammonia cools and condenses and, through

capillary action, travels back to the heat source under the electronics. We see the mass

and volume breakdown of the thermal control components for the electronics in Table

5.3.6-1.

Xenon Tank Thermal Control

To keep the xenon within its permissible storage temperatures, we embed a 5-Watt

resistance heating wire in the tank with thermostat regulation. We also wrap the tank

with Multi-Layer Insulation in order to mitigate heat loss by radiation.

Electric Thruster Thermal Control

The electric thruster is 55% efficient at converting its input power to thrust. The thruster,

therefore, generates approximately 639W of heat that we need to dissipate to keep the

thruster under its maximum operating temperature of 473K (200ºC). To maximize the

emissivity of the thruster, we paint the thruster with Z93 white paint. This increases the

amount of heat that the thruster is capable of radiating to space. At the base of the OTV,

a magnesium alloy shroud is installed around the thruster. The surface area of this shroud

is more than adequate to sufficiently radiate the thruster‟s waste power to space.



We see the dimensions of the OTV‟s thermal control system in Table 5.3.6-1. Because

we account for the aluminum shroud in the structure group‟s mass budget, we do not

consider the shroud in this table.

Table 5.3.6-1 100g Payload OTV Thermal Control Dimensions

System Component Mass (kg) Dimensions

Electronics Ammonia ~ 0 -

Heat Pipes 2.22 I.D. = 3.35cm; O.D. = 3.65cm;

Length = 5m

Radiators 1.21 Total Cross-Sectional Area =

0.895m2

(8 fins @ 0.112m2 each)

Xenon Tank Wire Heater 1 -

Electric Thruster Aluminum Shroud N/A N/A



5.3.7 – Power

Introduction

The operational life of the Orbital Transfer Vehicle (OTV) extends to at least one year.

We employ a light-weight and reliable power generation system to provide sufficient

power for this time span. The OTV power is provided by two circular Ultraflex solar

arrays and a secondary (rechargeable) lithium-ion battery. During periods in sunlight, the

arrays provide adequate power to meet the OTV‟s requirements. We do not assume,

however, that the spacecraft remains in sunlight throughout the transfer. The

rechargeable battery provides power during periods of darkness. Figure 5.3.7-1

illustrates a CATIA computer model of one of the partially deployed solar arrays.

Fig. 5.3.7-1 Top and side views of OTV Ultra-flex solar array CATIA model; partially deployed.

(Ian Meginnis)



Power Budget

We see the power budget for the OTV during the translunar phase broken down by group

in Table 5.3.7-1. The largest contributing factor to the OTV power budget arises from

the electric thruster. This device consumes over 75% of the entire OTV power budget.

To ensure all of the power needs are adequately met, we increase the total power

production by 5%. This increase accounts for fluctuations in power requirements by each

group that occur throughout the mission. The specific components for the individual

systems are mentioned in Section A-5.3.7 of this report.

Table 5.3.7-1 100g Payload OTV Power Budget by Group

Group Power Units

Propulsion 1529 Watts

Communication 0 Watts

Attitude 101.4 Watts

Power 120 Watts

Lunar Lander

(during translunar) 105 Watts

TOTAL 1959 Watts

Solar Array Sizing

With the power budget, we determine the size of the solar arrays. The arrays are sized to

meet the power requirements of the OTV during sunlight. Able Engineering produces the

two Ultraflex-175 solar arrays, which are comprised of triple-junction gallium-arsenide

solar cells. The power density for the arrays dictates the required mass and area. We see

the mass, stowage volume, and deployed area for the two solar arrays in Table 5.3.7-2.

Table 5.3.7-2 100g Payload OTV Solar Array Dimensions (total)

Parameter Value Units

Mass 13.06 kg

Stowage Volume 0.0664 m3

Deployed Area 6.54 m2



During launch, we store the circular solar arrays in two rectangular boxes on opposing

sides of the OTV. After the OTV reaches Earth parking orbit, we unfurl and deploy the

arrays. We implement a sun sensor and two motors to ensure the arrays provide

maximum power. The motors provide two degrees of freedom for each array.

The high cost of the solar arrays directly affects the orbit trajectory and engine selection

for the OTV. We optimize the vehicle design to reduce overall mission cost. The cost

per watt of the triple-junction solar arrays is approximately $1000, which results in a total

cost of $1.96 million.

Battery Sizing

With the power budget, we also determine the size of the battery. The battery is sized to

provide power to the OTV during periods of darkness and prior to solar array

deployment. The rechargeable lithium-ion battery, manufactured by Yardney, has a

certain power density that determines the battery dimensions. We see these dimensions

in Table 5.3.7-3.

Table 5.3.7-3 100g Payload OTV Battery Dimensions


Mass (includes housing) 12.17 kg

Volume 0.0043 m3

Total Energy 1534 W-hr

Power Subsystem Components

The power system for the OTV includes all devices necessary to generate, manage, and

distribute power to the individual OTV components. These devices include:

Solar Arrays

Batteries

Power Conditioning and Distribution Unit (PCDU)



Power Processing Unit (PPU)

Direct Current / Direct Current (DC/DC) Converters

Fig. 5.3.7-2 OTV Subsystem Power Description.

(Ian Meginnis)

The solar arrays generate approximately 1.96kW with a direct-current output voltage of

~200V. This power first travels to the PCDU. Broad Reach Engineering custom

fabricates the PCDU. The PCDU conditions and regulates power from the solar arrays

and distributes this power to the battery, PPU, and DC/DC converters.

The OTV‟s battery is available as a commercial off-the-shelf product. The battery

consists of ten 3.6V, 43 A-hr lithium-ion battery cells. These cells group together to

produce the required 1.53kW-hr of energy that we need for the battery. If the battery is

not fully charged, we route excess power from the solar arrays to the batteries via the

PCDU. During periods of darkness, the PCDU detects a drop in power from the solar

arrays and draws power from the batteries to support the OTV components.

Not to

scale



The PPU also receives power from the PCDU. The PPU increases the voltage from 100V

to the 340V the electric thruster requires. In order to power the OTV‟s individual

components, we use a series of DC/DC converters. These devices decrease the voltage

that comes from the PCDU to whatever voltage is required by the individual components.

Because each component has a unique input voltage requirement, a separate DC/DC

converter exists for each component. A total of six DC/DC converters are used to

support the individual OTV components. Modular Devices Incorporated produces the

converters, which are available as off-the-shelf products.

Table 5.3.7-4 100g Payload OTV Power Subsystem Dimensions

Component Mass (kg) Volume (m3)

Solar Arrays 13.06 0.0664 (stowed)

Battery 12.17 0.0043

PCDU 8.88 0.0157

DC/DC Converters 1 0.000218

Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1,

Author: Ryan Nelson

5.4 – Lunar Descent

5.4.1 – Structures

Lander Structure

The most important consideration in designing the structure of our Lunar Lander for the

100g payload is the mass of the frame. We identify two key drivers for the mass of this

frame. These drivers are total volume and total mass of the Lunar Lander at lunar

touchdown.

The shape of the Lunar Lander for the 100g payload case is a conic frustum. A conic

frustum geometry can be pictured as a large cone with the top part chopped off as

depicted in Fig. 5.4.1-1. There are two reasons for using a conic frustum frame design.

First, a conic frustum houses the subcomponents of the Lunar Lander (such as space

balls, H202 tank, and main engine) while minimizing volume. Minimization of volume

also lowers the mass for thermal control. The various frame component masses can also

be recalculated easily for a conic frustum design without a Finite Element Analysis

(FEA) computational package. We provide the overall dimensions of the frame needed to

house all of the Lunar Lander subcomponents in Table 5.4.1-1; the total volume of the

Lunar Lander is 1.05 m3.

Table 5.4.1-1 Basic Frame Dimensions for 100g payload Lunar Lander


Height 1.0 meters

Bottom Diameter 1.3 meters

Top Diameter 1.0 meters

Length of Legs 0.607 meters

Figure 5.4.1-1 illustrates the main parts of the frame. The frame floor consists of a

circular outer ring and circular inner ring connected by four rectangular floor support

beams. An additional circular ring makes up the top of the Lunar Lander. The top

circular ring connects to the frame floor by four hollow circular side support beams.


Author: Ryan Nelson

Fig. 5.4.1-1 Listing of Basic Lunar Lander Frame Components.

(Ryan Nelson)

These side support beams support the various loads subjected to the Lunar Lander

throughout the mission while maintaining a low thickness. The low thickness of the side

supports enables the storage of the four Lunar Lander legs within these side supports

during Earth launch and Lunar Transfer. A 0.5mm magnesium skin placed around the

entire Lunar Lander frame protects against micrometeorites and provides thermal

protection. The sum of all these frame components yields a total mass for the Lunar

Lander frame of 11.44 kg.

The total mass of the Lunar Lander at lunar touchdown determines the thickness, size and

mass of the individual frame components. We use a safety factor of 1.5 for all frame

components in the structural sizing code, Lander_Frame_ball.m. With the dry Lunar

Lander mass at touchdown and the Earth g‟s experienced at lunar touchdown, the forces

seen on the Lunar Lander frame are calculated. The size and mass of the different frame

components are then calculated based on the types of loading that cause initial failure to

get the overall frame mass.


Author: Korey Lemond

Material Selection

The lander is made of the same materials that the OTV is fabricated from. The floor and

shear panels are fabricated from magnesium AZ31, while all other structural members are

made of aluminum 6065. All bolts, rivets, fasteners, and connectors are made of AIS

1015 low carbon steel.



CAD

We see in the following figures a CAD model created in the program CATIA. This model

gives us a visual representation of what our Lander looks like, and gives us a general idea

of where certain systems are located. The Lander assembly is designed with a color

coding system in mind to easily identify certain subsystems.

1) Red – Structures

2) Orange – Attitude

3) Yellow – Power

4) Green – Communication

5) Blue – Propulsion

6) Tan – Space Ball Housing



Fig. 5.4.1-2 External View of the Lander with legs deployed.

(Joshua Elmshaeuser)

1 meter

H2O2 tank



Fig. 5.4.1-2 Close up of the Landers internal components with legs stowed.




Fig. 5.4.1-3 View of the Landers engine systems.


1.3 meters



Fig. 5.4.1-4 Visualization of the Lander on the Moon.

Earthrise background is public domain on NASA’s Earth Observatory website:

http://earthobservatory.nasa.gov/IOTD/view.php?id=4882



Author: John Aitchison


Introduction

The purpose of this section of is to provide a complete description of the lunar descent

trajectory and landing site. We provide an overview of the descent phase beginning from

lunar parking orbit, as well as a description of the landing site. See Section 5.3.2 for a

description of the flight trajectory up to lunar parking orbit injection.

Descent Overview

At this stage, the Lunar Lander has separated from the orbital transfer vehicle and is in a

25 km circular orbit around the Moon. This orbit will be referred to as the lunar parking

orbit. The spacecraft remains in this orbit while descent system checks are performed.

After completing these checks, a command is sent to the Lunar Lander to begin descent.

A small burn is performed by the attitude thrusters, which places the Lunar Lander in an

elliptical trajectory with a perilune altitude of 13.4 km. This elliptical orbit is referred to

as the lunar descent transfer orbit. When the Lunar Lander reaches perilune of this

elliptical orbit, we begin final descent with Lunar Lander main engine ignition. Final

descent lasts 207 seconds culminating in touchdown on the surface of the Moon. Figure

5.4.2-1 gives a pictorial overview of the descent.

Fig. 5.4.2-1. An overview of the entire lunar descent. Note: Not to scale (John Aitchison)

Lunar Parking Orbit

Lunar Descent Transfer Orbit

Final Descent

Moon



Final descent is a continuous burn with four distinct phases: radial, rotation, vertical, and

landing coast. Figure 5.4.2-2 gives a pictorial overview of the descent. Please note that

the shape of this trajectory is not to scale; for to scale depictions of the trajectories see

section A-5.4.2.

1.) Begin radial burn

2.) End radial burn, begin rotation

3.) End rotation, begin vertical burn

4.) End vertical burn, coast to landing

Fig. 5.4.2-2 An overview of the final descent. After radial burn, the Lunar Lander throttles down

and rotates to vertical, then throttles up to maximum and begins vertical burn and coasts to a

landing. Note: Not to scale

(John Aitchison)

Tables 5.4.2-1 and 5.4.2-2 provide high level overviews of the descent.

Table 5.4.2-1 Final Descent Action Breakdown

Action Time (s) Thrust (N)

Radial Burn 000 - 176 1150

Rotation 176 - 181 115

Vertical 181 - 196 1150

Coast to Land 196 - 207 130

2

3

1

4



Table 5.4.2-2 Descent Propellant Masses

Action Time (s) Propellant

Mass (kg)

De-Circularize

0.4

Radial Burn 000 - 176 67.5

Rotation 176 - 181 0.2

Vertical 181 - 196 5.8

Coast to Land 196 - 207 0.5

Unusable

2.3

Contingency

1.5

Total

78.2

Footnotes: This is the total propellant mass in

LPO.

Landing Site Approach

As the Lunar Lander approaches the landing site on final descent, it is critical that

adequate clearance between the spacecraft and mountainous terrain of the moon be

maintained. At all times, a clearance of over 6 km is maintained between the Lunar

Lander and the spacecraft, as depicted in Fig. 5.4.2-3.

Fig. 5.4.2-3 The green line is the descent trajectory, the blue is representative mountainous terrain

on the Moon. At all times, more than 6 km of clearance is maintained. Note: Not to scale (John Aitchison)



Landing Location

We choose Mare Cognitum (4°S 23°W) as our landing site. Figure 5.4.2-4 shows the

Moon and the location of Mare Cognitum. The landing location is relatively flat for

approximately 20 km in all directions. Immediately beyond this area there are

mountainous regions with altitudes of 2780 m. At a distance of 60 km from the central

landing location the altitudes reach a maximum of 4300 m.

Fig. 5.4.2-4 Landing site Mare Cognitum.

(www.science.nasa.gov)

Located about 20 km northwest from the landing site is the original landing of the Apollo

12 mission. Our vehicle is equipped with a camera to capture an image and video of the

landing location, where artifacts from the mission were left behind. Recording this

historical site will fulfill the Google Lunar X PRIZE Heritage Site Bonus. Figure 5.4.2-5

shows the anticipated landing site, the landing site of Apollo 12 and the surrounding

terrain.

Mare Cognitum



Fig. 5.4.2-5 Topographical map of the landing site.

(NASA/USGS/LPI/ASU)


Author: Thaddaeus Halsmer

5.4.3 – Propulsion

Overview

We chose a radial flow hybrid engine as the Lunar Lander main engine. The oxidizer is

98% hydrogen peroxide (H2O2) and the solid fuel grains are Polyethylene (PE). As

shown in Fig. 5.4.3-1, we inject H2O2 around the perimeter of the combustion chamber

between the PE fuel grain plates. The combustion gases exit the combustion chamber

through the center hole in the lower fuel grain plates and into the nozzle. Catalyzing

chemicals integrated into the PE fuel grains causes rapid decomposition of the H2O2 and

allows raw H2O2 to be injected directly into the combustion chamber.

Fig. 5.4.3-1 Radial flow hybrid engine dimensions and configuration

(Thaddaeus Halsmer)

The dimensions of the combustion chamber are dictated by the regression behavior of the

fuel grain plates, total burn time and the desired oxidizer to fuel mixture ratio. We size

the chamber based on these properties and the desired thrust of the engine. The resulting

dimensions are shown in Fig. 5.4.3-1 along with the exit diameter of the nozzle.



Performance/Operating Parameters

The Lunar Lander propulsion system is capable of being throttled to 10% of its maximum

thrust by adjusting the mass flow rate of the H2O2. Engine throttling allows us to meet

the retro-burn thrust and relatively low touchdown thrust requirements with a single

engine. Table 5.4.3-1 summarizes the primary performance specifications for the Lunar

Lander propulsion system.

Table 5.4.3-1 Hybrid Engine Performance Specifications

Parameter Specification Units

Thrust Max / Min 1100 / 110 [N]

Isp, vacuum average 320 [s]

Chamber Pressure Max / Min 2.1 / 0.21 [MPa]

Nozzle Area ratio Ae/At 100 ––

Burn Time 198.6 [s]

Propellant Feed System

We selected a stored gas–pressurization engine feed system with helium being used as

the pressurizing gas. Figure 5.4.3-2 is the feed–system fluid schematic showing the

architecture of the fluid systems. Also shown is the helium purge system that clears the

oxidizer feed lines of any foreign debris before engine firing and flushes out residual

oxidizer upon shutdown. Table 5.4.3-2 presents the description of the corresponding

fluid control components.

Table 5.4.3-2 Fluid control components

Item Description

SV01 Ullage Solenoid Valve

SV02 Purge Solenoid Valve

MOV Oxidizer flow control valve

CK01 Check valve

CK02 Check valve

HV01 Manual vent/fill valve

HV02 Manual drain/fill valve

REG Ullage pressure regulator

F01 Oxidizer filter/screen



Fig. 5.4.3–2 Lunar Descent Propulsion Fluid Schematic.

(Thaddaeus Halsmer)

Power Requirements (during operation)

The number of electronic fluid control components in use at any given time dictates

electrical power requirement of the propulsion system during operation. Based on the

known power requirements of the individual components, the power required for the

propulsion system will not exceed 124 watts during its use.


Author: Saad Tanvir

H2O2 Storage

The hydrogen peroxide is stored in a spherical aluminum tank. We use Aluminum 1060

due to its ability to reduce the Active Oxygen Loss (AOL) from the H2O2 during the

mission. For a sealed Aluminum 1060 container the AOL for 90% H2O2 for 1 year is

0.1% of the initial hydrogen peroxide amount (Ventura, 2005). Hence, the amount of

H2O2 decomposed over the duration of 1 year for our mission is 0.078 kg. As compared

to the 78 kg of total propellant carried, this value is negligible.

The hydrogen peroxide tank is sized based on the tank operating pressure and the amount

of H2O2 used during the mission. Table 5.4.3-3 describes the pressure, size and mass of

the H2O2 tank. We use a safety factor of 1.5 in determining the following parameters.

Table 5.4.3-3 Pressure, Size and Mass of the H2O2 tanks


Operating Pressure 3.07 MPa

Volume 0.0533 m3

Thickness 1.075 mm

Diameter 0.467 m

Mass 2.06 kg

The total volume of the oxidizer tank is the sum of the volume of the H2O2 required for

the mission (0.0512 m3: including the usable propellant volume, boil-off volume and the

volume of unusable propellant) and a 4% ullage volume (0.00205 m3). Ullage is the

additional volume added to the tank to give the propellant space to expand in case of

increasing temperature and varying pressures inside the tank.

Pressure Feed System

Helium is used as the pressurant gas for the pressure-fed system. We chose Titanium as

the material for the He tanks. Table 5.4.3-4 shows the detailed specifications of the


Author: Saad Tanvir

pressurant tank as well as the mass of the Helium gas required to complete the landing

mission.

Table 5.4.3-4 Pressurant System parameters


Operating Pressure 21 MPa

Volume 0.0257 m3

Thickness 5.76 mm

Diameter 0.366 m

Mass of Tank 0.87 kg

Mass of Helium 0.89 kg

Propulsion System Inert Mass

The inert mass of the propulsion system is largely dependent on the amount of propellant

used during lunar descent. With propellant masses obtained from the trajectory analysis

done by mission operations, we size the chamber, the oxidizer (H2O2) tank and the

pressurant system. We use historical as well empirical relations to size the rest of the

propulsion system. The total propulsion system inert mass for the hybrid landing engine

is approximately 30 kg. Table 5.4.3-5 gives a detailed breakdown of the propulsion

system inert masses.

Table 5.4.3-5 Propulsion system inert mass breakdown


Oxidizer (H2O2) Tank 2.06 kg

Pressurant Gas (He) 0.89 kg

Pressurant Tank (He) 0.87 kg

Nozzle 12.28 kg

Combustion Chamber 3.15 kg

Injector 0.95 kg

Skirts and Bosses 0.32 kg

Feed System 0.32 kg

Valves 6.94 kg

Structural Supports 2.15 kg

Propulsion System Inert Mass 29.91 kg


Author: Saad Tanvir

Thermal Analysis

The operating temperature of the propulsion system during the lunar transfer phase is

decided based on the safe operating temperature range of the H2O2 and the polyethylene.

We chose an operating temperature of 10 o

C (283 K). The tanks are insulated with Multi-

Layer Insulation (MLI). We require 35W of constant power to maintain this temperature

throughout Lunar Transfer.

During lunar descent, the temperature of the propulsion system drops. The temperature

drop in the hydrogen peroxide tank is 6.0 K (from 283K to 277K). Similar temperature

drop from the valves and the feed lines are 2.0K and 2.1K respectively. Because the drop

in temperature of these components during lunar descent is small, its affect on engine

performance is negligible. Hence, there is no power required for thermal control on the

propulsion system during lunar descent.


Author: Christine Troy

5.4.4 – Attitude Control

During lunar descent, we use the hybrid propulsion descent engine to slow the Lunar

Lander to a safe landing speed. Ideally the thrust points directly through the vehicle‟s

center of mass. In practice, the thrust is offset from the center of mass due to (for

example) manufacturing limitations and center of mass changes as propellant is burned.

In addition to the thrust-offset torque acting on the vehicle, there are also much smaller

environmental perturbing forces that act on the Lunar Lander. We discuss the effect of

these perturbing forces in greater detail in Section 5.3.4 (which pertains to attitude

control during Lunar Transfer). During lunar descent, the environmental forces are so

much smaller than the thrust-offset force that we neglect them in our analysis.

The sensors of our attitude control system are described in Section 5.3.4. We place these

sensors on the lander so that they may be used in both Lunar Transfer and lunar descent.

A laser altimeter is also placed on the Lunar Lander to find the distance to the surface

during descent.

The attitude control actuators on the lander consist of twelve hydrogen peroxide thrusters.

We choose hydrogen peroxide thrusters because they draw their propellant from the

descent-engine propellant tank and thus save on complexity and cost. These thrusters

(manufactured by General Kinetics) are each capable of producing 13.3 N of thrust. The

twelve thrusters are arrayed around the vehicle as shown in Fig. 5.4.4-1. This

configuration provides attitude control around all three of the vehicle‟s axes.

Additionally, the attitude control thrusters provide the small initial burn beginning the

descent phase.



Fig. 5.4.4-1 Lander thruster configuration schematic for 100g payload case.

(Christine Troy)

The properties of the vehicle and the thrusters govern the amount of attitude control

propellant required. The two most crucial factors in the amount of propellant required

are the distance from the vehicle‟s center of mass to the thruster and the specific impulse

of the thruster. We select the vehicle configuration and the attitude thrusters and then we

can calculate how much propellant we need. Table 5.4.4-1 summarizes the relevant

properties of the lunar descent attitude control system, including the amount of propellant

needed for attitude control.

Table 5.4.4-2 Properties of lunar descent attitude control system for 100g payload


Torque about each axis 25.2 Nm

Thrust available for lunar descent 56 N

Attitude control propellant mass 3.21 kg

Total mass of attitude thrusters 1.08 kg

Lander Top View

Lander Side view




Communication Hardware/Equipment

Two communication systems are installed onboard the Lander in order to communicate

with Earth and our Space Balls. One system communicates with Earth. The other system

communicates directly with the Space Balls. These systems operate at a data rate of 51.2

kbps at a frequency of 2.2 GHz. The total mass, power usage, and purchase price is listed

in Table 5.4.5-1.

Table 5.4.5-1 Communication Equipment Totals

Mass (kg)

Power Usage (W)

Purchase Price (2009 $)

2.48

60.53

315,668

The communication systems consist of:

1) Two antennae to communicate with Earth.

2) Two antennae mount pivots.

3) A transmitter to transmit to Earth.

4) A receiver to receive signals from Earth

5) An antenna to communicate with the Space Balls.

6) A transceiver to transmit to and receive signals from the Space Balls.

7) A video camcorder to record video and take pictures

8) A Computer board to control these systems


Author: John Dixon

Communication Power Requirements

The Lander sends information back to an Earth station at a data rate of 51.2 kbps. The

signal from the Lander requires a minimum of 33 Watts to get enough information

packets to the ground station. The frequency of the signal is 2.2 GHz, and allows

constant communication from the lunar surface.



Communication Timeline (and completion of GLXP requirements)

During the descent from lunar orbit to the surface of the moon our Lunar Lander is still

engaged in the 0.5 kbps (kilobytes per second) data link. This link is a continuation of the

data link used during the Lunar Transfer phase of the mission. This link still includes

telemetry from the spacecraft to Earth, and commands from Earth to the spacecraft.

Once on the lunar surface we begin our required transmissions to complete the Google

Lunar X PRIZE (GLXP) mission requirements. These files that we transmit to meet

mission requirements are outlined in Table 5.4.5-2. Photo files are produced from the

onboard digital camera. X PRIZE Foundation (XPF) set asides are files that are provided

by XPF and are stored on the craft‟s computer memory before launch. These XPF set

asides include the first e-mail and text messages to be sent from the Moon. The data

uplink set is provided by the XPF, transmitted to the Moon, and then downloaded back to

the Earth. After the Locomotion phase of the mission begins, more mission requirements

will be transmitted. These transmissions are discussed in the Locomotion section of this

report, Section 5.5.3.

Table 5.4.5-2 Lunar Arrival Mooncast

Item Link Directiona

Size [MB] Transmission Time [hr]

Photos Down 5 0.24

XPFb Set Asides Down 10 0.47

Data Uplink Set Up 10 0.47

Data Uplink Set Down 10 0.47

Footnotes:

a – Down is from Moon to Earth; up is from Earth to Moon.

b – X PRIZE Foundation.


Author: Kelly Leffel


Lunar Day

We need to maintain operable temperature of the Lunar Lander from Lunar Transfer

through the lunar night. During Lunar Transfer, the side of the Lunar Lander that faces

the sun is at 394 Kelvin while the side out of the sun can be 2 Kelvin (Scott, 2009). The

temperature gain from the sun needs to be transported through the Lunar Lander to the

shaded side. During a portion of Lunar Transfer, part of the mission is spent where the

earth completely shades the sun from the Lunar Lander. During this time the Lunar

Lander‟s internal temperature decreases. The Lunar Lander requires different types of

thermal control to maintain the operable temperature.

We use a multilayer insulation (MLI) blanket to reduce the amount of heat absorbed by

the Lunar Lander and to keep the heat in during the lunar night. As the surface area of the

Lunar Lander increases, the heat absorbed during the day increases and the heat out

during the night increases. For the 100g payload case, the MLI blanket will cover the

entire outside of the Lunar Lander, the Space Ball compartments, and the propulsion

system. The propulsion system and the Space Ball compartments will need the MLI

blanket because these systems should not undergo temperature differences.

We use a passive cooling device to reject heat in the Lunar Lander. This thermal control

system is similar to the thermal control for the orbital transfer vehicle (OTV). This

system includes an aluminum heat plate, a heat pipe and a radiator. The aluminum plate

is used to conduct the heat from the communication equipment. The communication

equipment sits on this aluminum plate. The area of the heat plate is the sum of the bottom

surface area of the communication equipment. This system can be seen in Fig. 5.4.6-1.

The ammonia heat pipe runs underneath the heat plate. This pipe is five feet long and

connects the aluminum plate to the radiators. Increasing temperature in the Lunar Lander

vaporizes ammonia in the heat pipes. Ammonia vapor flows to radiators via capillary



action, rejecting heat from the Lunar Lander system. The ammonia will liquefy and be

brought back through the ammonia pipe by capillary action. This action is continuous

during the mission.

Fig. 5.4.6-1 Schematic of Passive Thermal Cooling System.

(Josh Elmshaeuser)

The heat pipe connects to the radiator on the Lunar Lander. The radiator on the side that

faces the sun allows the ammonia to flow through the radiator with little or no

temperature increase. The radiators are painted white to reflect this heat. The radiator in

the shade of the Lunar Lander rejects the heat flowing through the heat pipe. Figure

5.4.6-2 shows this heat exchange. The radiator‟s area is large enough to expel the heat the

Lunar Lander gains at a peak heat time. The peak heat time is on the lunar surface when



the sun covers the entire Lunar Lander and all the communication equipment is running.

The radiator consists of aluminum and sits at an angle where the radiator will be able to

expel the heat from the bottom when the sun is directly overhead.

Fig. 5.4.6-2 Heat transfer in the Lunar Lander - Heat enters on the side with the sun and

leaves on the shaded side both through the Lunar Lander and out the radiators.

(Kelly Leffel)

We operate two valves, one on each heat pipe system, to control the temperature of the

Lunar Lander. These valves are the yellow section on Fig. 5.4.6-2. As the temperature of

the Lunar Lander decreases due to the lack of sun, either in the earth‟s shade or the lunar

night, the valves will close. This process eliminates the rejection of heat when the Lunar

Lander is at the minimum operable temperature.

We use heaters to control the temperature of some components of the Lunar Lander

during Lunar Transfer. The propellant and oxidizer tanks along with the ball

compartments use resistance heaters to replace heat lost to space throughout the Lunar

Transfer. The resistance heaters replace the amount of the heat lost by the equipment.

Two additional heaters sit on the heat pipe where the valves are located. These heaters



provide heat to the ammonia if the temperature of the radiator drops below 195 Kelvin,

the freezing temperature of ammonia (Gilmore, 2002). The total power required by the

thermal control system during Lunar Transfer is 10 Watts.

The mass of the thermal equipment was a significant driver in choosing a thermal control

system. The current system that we use has a total mass of 9.57 kilograms. The

breakdown of the thermal control subsystems are found in Table 5.4.6-1. The total mass

was calculated in the MATLAB program LanderThermalControl.m.

Table 5.4.6-1: Day Thermal Control of Lunar Lander


MLI Blanket 2.35

Aluminum Plate 1.40

Heat Pipe 2.60

Radiators 2.70

Ammonia 0.021

Heaters 0.50

Total 9.57


Author: Tony Cofer and Adham Fakhry

Surviving the Lunar Night

The ambient temperature of the lunar surface during its fourteen day dark phase

plummets to 140 degrees Kelvin which is far below the tolerance ranges for many of the

Lunar Landers‟ components. Heat is lost through conduction by contact with the surface

and through radiation. We minimize thermal losses by using multi-layer insulated

blankets to reduce radiation losses from the protected volume. Total thermal losses for

the 100 gram payload configuration are expected to be about 11 Watts which must be

replaced during the period of darkness.

We use chemical energy to provide heat to the Lander during the lunar night. We choose

hydrazine for its simplicity of use and for its high energy content. Less than four

kilograms of hydrazine are required to survive the lunar night, and is contained in a small

tank wherein a rod of Aluminum/Nickel catalyst is extruded by a bimetallic actuator as

seen in Fig. 5.4.6-3.

Fig. 5.4.6-3 Hydrazine heating system (right) and catalyst actuator (left).

(Tony Cofer)


Author: Tony Cofer and Adham Fakhry

The thermostatic element must be experimentally calibrated to provide the desired

equilibrium temperatures throughout the thermally controlled volume

We minimize power requirements during lunar night by shutting down all electrical

equipment. A small electrical switch powers the Lunar Lander down when solar power

drops below a predetermined threshold and turns it on again at the lunar dawn.


Author: Adham Fahkry

5.4.7 – Power

Battery

We size the battery for the Lander based on power requirements from the Lander engine,

communication and attitude systems. The Lander engine, the communication gear and the

attitude systems all depend on the battery during the landing phase.

The Lander‟s engine requires power to operate the various valves and pumps to make the

engine function, which total 124 Watts and will be operating for 250 seconds. The

communication gear uses a peak total power of 60.53 Watts, to ensure that we can

maintain constant communication between the Earth and the Lander throughout the entire

Landing phase. We provide enough power for the communication gear to transmit for 27

minutes at peak power consumption. The attitude system uses a peak total power of 25.4

Watts to power our system and requires 15 minutes of peak power to make sure the

Lander arrives at its destination. Table 5.4.7-1 highlights the total power requirement for

all systems and their individual components on the 100 g Lander.

Note: Complete description of each device can be found

in section 5.4.2

For a complete background on each device within each system, please refer Section A-

5.4.7.

We purchase the batteries from the company Yardney Technical Products, Inc., who

provide readymade lithium ion batteries that are used in space vehicles.

Table 5.4.7-1 Power breakdown of system on Lander

Item Power Units


Communication 60.53 Watts

Attitude 25.4 Watts

Total 209.93 Watts



We select a 12 Ampere-hour battery, but the battery only requires 11.69 Ah. We decided

that as a safety precaution the battery should have a little bit extra power, in case of any

problems that might occur after separating from the OTV. Refer to Section A-5.4.7 to see

how we calculated the capacity of the battery in Ah.

In order to avoid loss of charge when not in use, which happens with all lithium ion

batteries, our battery on the Lander will be completely drained during integration with the

entire spacecraft. The team will send a command that will allow the battery to charge

when it approaches the moon, using the OTV solar cells before separation. We compute

the battery size using lander100g.m.

Solar Array Sizing

We use solar cells to provide enough power to the communication system during the two-

week long lunar day. The peak power required is 60.53 W to run all of the

communication equipment, refer to Section A-5.4.7 to see breakdown of power

requirements for each device within the communication equipment.

Based on these specifications, we set up the solar cells in a deployed configuration for the

entire mission. The cells are statically fixed to reduce costs and complexity. The cell has

an area of 0.785 m2. Figure 5.4.7-1 displays the power consumed and the power

generated by the cells through a lunar day.



Fig. 5.4.7-1 Maximum Potential Power of solar cells during a lunar Day.

The communication gear has enough power to meet its peak power requirements

on the 1.7st day of the lunar day. At maximum power, the solar cells provide

74.5% more power than required at day 7 of the lunar day.

(Adham Fakhry)

The red line represents the power used by the Communication gear, the blue line

represents the power produced by the cells. The solar cells provide more than enough

power required by the communication gear. Also, the excess power ensures the dust

accumulated on the lunar surface or impacts from meteorites, will not severely affect the

power collected by the cells as they provide extra power. We sized the solar cells to meet

the communication system‟s power requirements during data transmission from the Earth

to the Moon.

The solar cells are provided by Able Engineering, who will be supplying solar cells for

the Orbital Transfer Vehicle and the Lander. The Lander solar cells will cost $250,000

and the solar cells will also weigh a maximum of 2 kilograms. Solar array sizing is



automated in lander.m. Refer to Section A-5.4.7 for breakdown of costs and the

constants used to size the solar cells.

DC-DC Converters

The DC-DC converters for the 100 gram case depend on the number of component within

the system and the voltage they each require to operate, the DC-DC converters change

the voltage from the battery and the solar cells to a higher voltage for each individual

device will require. Table 5.4.7-2 below displays the DC-DC converter system for the

100 g payload.

Table 5.4.7-2 DC-DC converter system

Component Quantity Units

Mass 0.725 kg

Dimensions 0.033 x 0.033 x 0.033 m

Temperature Range -50 to 150 C

Cost $ 21,000 - Note: Cost is in 2009 US dollars and is the cost for all the units.

All the converters will be placed in a single aluminum box that will be connected to the

PCDU (power conditioning and distribution unit) that will then branch out to each system

and their various components.


A power conditioning and distribution unit (PCDU) is needed to safely manage the power

distribution throughout the Lander during mission. Refer to Section A-5.4.7 for complete

breakdown of the PCDU. We choose a PCDU from Terma Incorporated to handle a 240

Watts power system that will be connected to our power system components, listed

below, and the PCDU will cost $12,000 and weigh approximately 1.9 kg.

The PCDU will interface with the following sources:



1. Rechargeable Lithium-ion battery

2. Solar Cells

3. DC-DC converters

Table 5.4.7-3 shows the specifications of the PCDU.

Table 5.4.7-3 Specifications of the PCDU on the Lander


Mass 1.932 kg

Dimensions 0.033 x 0.033 x 0.033 m


Cost $12,000 - Notes: Cost is in 2009 US dollars.

Mission Configuration – 100g Payload – Locomotion Section 5.5.1,

Author: Cory Alban

5.5 – Locomotion

We have reached the moon. Everything goes silent as the engine shuts off. Suddenly, a

panel blows off of the bottom of the Lander and out pops a small round sphere the size of

a basketball. After a few moments, the ball begins rolling away from the Lander; quickly

picking up speed. Traversing the rolling lunar landscape and bouncing up and over small

bits of moon rock, the Space Ball is on its way to completing the mission we have come

so far to achieve.

Some basic characteristics of the Space Ball are listed in Table 5.5-1 but the following

sections describe all of the different systems that make up the locomotion phase of the

100g payload mission in greater detail.

Table 5.5 - 1 Space Ball Specifications

System Category Value Units

Mass 2.44 kg

Diameter 0.25 m

Cruise Speed 1.04 m/s

Min Turning Radius 0.0625 m

Fig. 5.5-1 Artistic Rendition of Space Ball on the Moon.

(Cory Alban)


Author: Ryan Lehto

5.5.1 – Structures/Integration

Space Ball Deployment

The Space Ball resides inside a thermally insulated box with protective foam surrounding

the lower half. We choose to deploy the Space Ball using a linear shaped charge. The

linear shaped charge is made of copper lining with C-4 as the explosive material. The

linear shaped charge surrounds the Space Ball on three edges of the bottom panel. The

fourth edge of the deployment door has a hinge. The charge dislodges the bottom of the

casing, which swings open on the hinge. The Space Ball then drops to the surface ready

to begin the mission. Figure 5.5.1-1 shows us a cross sectional view of the Space Ball

deployment system. Table 5.5.1-1 shows the mass of the deployment system.

Fig. 5.5.1-1 View of the Space Ball inside the thermally insulated box

surround by SOLIMIDE foam for protection from the linear shaped

charges that cut the bottom of the Lunar Lander for deployment.

(Caitlyn McKay)

Table 5.5.1-1 Mass of deployment system per Space Ball

System Mass (kg) per Space Ball

Charge 0.580

Foam 0.040

Total 0.620

Space Ball SOLIMIDE Foam

Linear shaped charge

Thermally

insulated box


Author: Ryan Lehto

CAD

We arrange the components as shown in Fig. 5.5.1-2. The outer shell seals the Space

Ball and the components inside. A main axle fixed to the shell houses the main drive

motor as well as a stepper motor for turning. In addition, the pendulum mass and the dust

mitigation motor mount to the drive axle. The 100 g payload, CPU, transceiver, camera,

and battery mount to the pendulum swing arms.

Fig. 5.5.1-2 Space Ball internal components.

(Ryan Lehto)

Transceiver

100g Payload

Camera

CPU

Main Axle &

Motor Housing

Dust

Removing

Vibrating

Motor

Outer Shell

0.25

m


Author: Ryan Lehto

Fig. 5.5.1-3 The Space Ball is

roughly the same size as a

basketball.

(Ryan Lehto)

An advantage to the Space Ball system is its small size. The Space Ball is similar in size

to a basketball (about 0.25 m diameter) see Fig. 5.5.1-3.



Structural Analysis

We choose Lexan, a clear polycarbonate material, for the shell of the Space Balls. The

clear outer shell allows us to mount our camera inside of the ball rather than having to

mount it to the exterior. Keeping everything inside of the ball reduces system complexity

and the number of failure modes. Every fragile component in the Space Ball is protected

by the sturdy Lexan shell. Due to the high impact resistance of Lexan, the Space Ball can

survive an impact with a rock at full cruising speed without cracking the shell or breaking

sensitive components. Pressurized nitrogen inside the Space Ball adds additional stress

to the shell.

The aluminum drive axle runs through the center of the ball and carries a bending

moment from the equipment mounted on the torsion arm. The axle has additional torsion

stress from the drive system.

Tables 5.5.1-2 and 5.5.1-3 describe both the physical structure of the Space Balls and

their load bearing capabilities.

Table 5.5.1-2 Space Ball Load Bearing Capabilities

Failure Mode Load Can structure handle load? Factor of Safety

Torsion Stress 0.03N-m Yes 59

Bending Moment 15.69N-m Yes 100

Pressure Vessel 101.325kPa Yes 2.8

Table 5.5.1-3 Space Ball Structure

Structure Component Size Shape

Lexan Shell 3.82mm Thick, 25cm Diameter Hollow Sphere

Aluminum Axle 25cm Long, 6cm Diameter Circular Rod

Aluminum Torsion Arm 12.5cm Long Rectangular Rod


Author: Cory Alban

Dust Mitigation

As the Space Ball rolls over the lunar surface, dust will accumulate on the shell of the

Ball. This gathering and sticking of the lunar dust can potentially cause problems. It can

block the view of the camera, prohibiting us from achieving the goals of the Google

Lunar X PRIZE. In addition, buildup of dust increases the risk of it entering the core of

the Space Ball where all the critical mechanical components are. The dust penetration

into the drive system could potentially wedge the mechanisms, causing a failure by

prohibiting them from moving.

Angstrom Aerospace has created inflatable robots to explore distant planets. They are

using an ultrasonic motor to vibrate off the dust (Marks, 2008). We have adapted their

method of dust removal by implementing a linear ultrasonic Piezo drive Motor to vibrate

the dust off the surface of the Ball. The model selected is M-674,164 PILine designed by

Physik Instrumente (PI). At only 38 mm2 and weighing 0.1 kg this option made it the

most efficient solution for dust removal. The motor creates a frequency of 155 kHz,

which vibrates the Space Ball. This frequency is more than sufficient to shake off the

dust.

The ultrasonic motor consists of a rod mounted between two piezo linear motors. The

rod of the motor is attached to the center shaft of the Space Ball so that both motors

oscillate back and forth at a high frequency. This intense motion causes the ball to

vibrate. The motor will not be continuously running it will only be used when dust

begins to accumulate. When the motor is needed the ball will not be in motion.


Author: Ryan Lehto


System Design

We implement a center of mass shifting system for propelling the Space Balls. The

Space Ball propulsion system works by swinging a mass, much like a pendulum to shift

the ball‟s center of mass. The system includes two electric motors, a continuous D/C

motor mounted on a main driveshaft, and a D/C stepper motor positioned perpendicular

to the main drive motor for turning. The motor system all together has a mass of 0.172

kg. A benefit of the space ball propulsion system is the use of the essential components

as part of the drive system. The pendulum mass includes the 100 g payload, CPU control

board, batteries, and the communication equipment.

Main Drive Motor

We use a 6 mm diameter D/C motor paired with a 6 mm planetary gear head as the main

drive system to propel the Space Ball for the mission. The gear multiplies the motor‟s

torque and slows down the motor‟s output shaft speed. The motor controls the

forward/back movement by lifting the hanging mass and thus shifting the Space Ball‟s

center of mass see Fig. 5.5.2-1. The motor gear combination creates a output torque of

0.03 N-m. The torque is more than capable of accelerating the space ball at 0.0043 m/s2

with an average velocity of 1.04 m/s. The Space Ball travels 500 meters in 8 minutes.

Fig. 5.5.2-1 The forward and backward motion of the

Space ball as controlled by the main drive motor.

(Ryan Lehto)


Author: Ryan Lehto

We control left to right movements with the stepper motor. The stepper motor steps and

holds the drive system pendulum mass in 15° increments up to 30° in the desired turning

direction. The left/right movement of the pendulum tilts main drive axle and as result

turns the Space Ball see Fig. 5.5.2-2. The space ball is able to turn in a radius of 0.0625

m.

Fig. 5.5.2-2 The pendulum moves left or right to turn the Space Ball.

(Ryan Lehto)

The main drive motor needs 0.5435 W to propel the ball. Thus, for the entire mission the

space ball requires 0.0725 W-hr to travel 500 m.



5.5.3 – Communications


The Lander relays communications from the Space Balls to Earth and vice versa. For

communications equipment on the Space Balls the limiting factors are size, and mass. We

require the smallest possible components that still have the capability to transmit back to

the Lander at a distance of 500 meters. During hardware selection we consider size,

power and price. The Space Balls contain the following four pieces of communication

equipment.

1) UHF Antenna

2) Transponder

3) CPU

4) Camera

In compliance with the GLXP guidelines any vehicle placed on the moon must take one

self-portrait during the mission, so it is necessary that we place a camera in the Space

Balls. We purchase the camera from Ahlberg Electronics. The camera is designed to be

radiation tolerant enough to handle nuclear power facilities and is suitable for subjection

to space conditions.

We will employ monopole UHF antennae and UHF transceivers which provide a high

data transmission rate over short distance. The UHF antennae allow quick, cheap, and

power efficient communication over the short distance that the Space Balls travel. The

Space Quest Company provides these required systems at the cheapest price.

We use a CubeSat Kit Flight Module as the CPU for the Space Balls. The CPU processes

all the information sent from the Lander to the Space Ball, as well as the pictures taken

by the camera on board.

Mission Configuration – 100g Payload –Locomotion Section 5.5.3,

Author: John Dixon


The communications equipment on the SpaceBalls was chosen after the power and

communications requirements were known. The SpaceBalls will send information back

to the Lander at a data rate of 51.2 kbps. The signal from the SpaceBalls requires a

minimum of 3 Watts to get enough information packets to the Lunar Lander. The

frequency of the signal is 2.2 GHz, and will allow constant communication from the

SpaceBalls furthest planned distance from the Lunar Lander.




After the Locomotion phase on the lunar surface we transmit proof of our locomotion to

fulfill Google Lunar X PRIZE (GLXP) mission requirements. These files consist of 8

minutes of real time video, 8 minutes of high definition video, and 5 megabytes (MB) of

photographs. The Locomotion Mooncast is outlined in Table 5.5.3-1.

Table 5.5.3-1 Locomotion Mooncast



8 min Near Real Time Video Down 75 3.53

8 min High Definition Video Down 900 42.37

Photos Down 5 0.24

Footnotes:

a – Down is from Moon to Earth; Up is from Earth to Moon

After our survival of the lunar night, for our survival bonus, we have yet again another

Mooncast. This Mooncast is named the Survival Mooncast, and it is outlined in Table

5.5.3-2. All of the same data is transmitted except for the high definition video.

Table 5.5.3-2 Survival Mooncast



8 min Near Real Time Video Down 75 3.53

Photos Down 5 0.24

Footnotes:

a – Down is from Moon to Earth; Up is from Earth to Moon

Mission Configuration – 100g Payload - Locomotion Section 5.5.4,

Author: John Dixon


The inside of the Space Balls will be filled with Nitrogen Gas. The Nitrogen Gas‟s

thermal properties enable it to absorb most of the heat present inside the Space Balls,

keeping the temperature rise to a minimum. The Aluminum frame inside the Space Balls

absorb the heat in the Nitrogen Gas and dispel it to space. The Nitrogen Gas has a very

low mass, and a high specific heat, making it an efficient, cost-mass effective cooler of

the Space Balls. The Aluminum frame is an essential support structure and is not

considered an additional mass in the thermal management system, despite acting as such.


Author: Jeff Knowlton

5.5.5 – Power

Power System

We select three Lithium Manganese Dioxide coin batteries to power each Space Ball.

The non-rechargeable primary batteries are provided by Efficient Energy & Managed

Battery. Each of these batteries provide 3 volts and 0.26 ampere-hours. The dimensions

for each of these batteries are 23mm in diameter and 3mm in height shown in Fig. 5.5.5-

1. This battery system produces 0.78watt-hours per battery and a total of 2.34 watt-hours

for all three batteries before the mission leaves Earth. After the twelve month transit to

the moon, the batteries lose 45% of the original charge.

Fig. 5.5.5-1 Lithium Manganese Battery CR2330 cylindrical

coin dimensions. [not to scale]

(Jeff Knowlton)

We chose these batteries based on a number of reasons, they provide enough power for

the mission and the correct voltage for components, the most predominate factor is that

they have a mass of only four grams which puts the total mass at twelve grams not

including housing. We budget out the power to each of the required systems as shown in

Fig. 5.5.5-2. The three batteries are sized to complete the 500-meter drive and transmit

data back to the Lunar Lander. The battery we are using provides a margin of safety of

1.52 for all of the power needs of the Space Ball.

3m

mm

23mm



Fig. 5.5.5-2 Space ball Battery distribution over

locomotion mission phase.

(Jeff Knowlton)

CPU Drive Motors

6%

Trans-mission

55%

Camera5%

Reserve34%

Battery Distribution


Author: Cory Alban


Lunar Surface Terrain

Lunar regolith is a fine and dry powder. The regolith has a small electric charge, which

causes it to stick to almost anything it touches. The adhesive dust coats the shell of the

Space Ball; however, we employ a dust mitigation technique to address this issue.

Table 5.5.6-1 shows the probability of encountering craters at our landing site.

Table 5.5.6-1 Lunar Crater Hazard Analysis

Crater Size Encounter

Probability

Is Space Ball

Capable?

Solution

Small (0-.02m) Very High Yes Roll Over

Medium (0.02m-4m) High Yes Avoid or Roll Through

Large (4m+) Very Low No Avoid

The Space Ball is large enough that craters of small size will not cause mobility

problems. The medium sized craters can be avoided by selecting an appropriate mission

path. The Space Ball can roll over and through medium sized craters while traveling at its

cruising speed. Large craters will be avoided.

Table 5.5.6-2 shows the probability of encountering rocks or other debris .

The majority of rocks the Space Ball encounters are of equal or smaller diameter as the

ball itself. In the event of a collision with a rock, the Space Ball is designed to withstand

a full speed impact with large objects. Just as a car wheel rolls over pebbles and small

rocks, the Space Ball runs up and over any objects that are smaller than the outer shell.

Turning capability and choosing an appropriate mission path will eliminate chances of

encountering large debris.

Table 5.5.6-2 Lunar Debris Hazard Analysis

Debris Size Encounter

Probability

Is Space Ball

Capable?

Solution

Small (0-.02m) High Yes Roll Over

Medium (0.02m-4m) Medium Yes Avoid or Dodge

Large (4m+) Very Low Yes Avoid


Author: Cory Alban

Completion of Mission Requirements

We complete the final mission requirements using the Space Ball. The following is a list

of our objectives.

1) The secondary vehicle deployed by the Lunar Lander must move a distance of

500m on the surface of the Moon in a deliberate manner.

2) The vehicle must carry a 100g payload to the surface of the Moon.

3) Transmit an “Arrival Mooncast” and a “Mission Complete Mooncast” to

Earth

To complete objective one, the Space Ball will complete a series of tasks on the lunar

surface. The Space Ball powers on and runs a systems diagnostic from inside of the

Lander to verify that all systems are working properly. Once all systems are up and

running, the ball is deployed onto the lunar surface. Mission control sends instructions to

the ball that indicate what direction to travel from the Lander based on the surrounding

terrain conditions. Photographs taken from the Lander shortly after touchdown and

satellite images of the area will be the primary factors in determining the navigation path.

The ball accelerates along the surface in the direction of travel until reaching 1.04 meters

per second. The Space Ball maintains this speed for eight minutes. At the end of eight

minutes of travel the ball reaches a distance of 500m from the landing vehicle and

achieves the mission objective.

To complete objective two, the thermal control, propulsion system, and structural

components of the Space Ball are designed with the knowledge that this 100g payload

will need to be carried the 500m distance and kept within reasonable thermal limits. By

satisfying objective one, objective two is subsequently met.

To complete objective three, the ball carries a camera and takes the required self portrait

picture and sends it back to the Lander. The ball turns itself ninety degrees to its right

during the braking maneuver at the 500m travel mark. This turning maneuver ensures that


Author: Cory Alban

the camera faces back toward the Lander. In the event that moon dust is obscuring the

camera‟s view, a vibration motor will run until the dust is shaken off and the view to the

Lander is clear. The ball transmits the image back to the Lander for eight minutes. Once

the Lander receives the entire image, it sends the data along with the other required

pieces of the “mission complete mooncast” back to Earth. The photo verifies that the ball

did travel the full distance and completes objective three.

Table 5.5.6-3 gives a breakdown of each step of the lunar mission along with a time

breakdown for the mission.

Table 5.5.6-3 Space Ball Mission Breakdown

Step Time(minutes) Tasks to be Completed

1 0 Space Ball performs a system diagnosis.

2 1 Deployment from Lander.

3 2 Direction of travel received from mission control.

Space Ball orients to path of travel.

4 2-10 Accelerate to cruising speed of 1.04m/s.

Travel for 8 minutes until 500m objective achieved.

5 11 Braking maneuver with a 90 degree orientation change to point

camera toward Lander. Shake off dust if necessary.

6 12 Snap photo of Lander from ball and begin transmission.

7 20 Finish Photo Transmission.

Mission Configuration – 100g Payload – Mission Integration Section 5.6,

Author: Korey LeMond

5.6 – Mission Integration

Two types of integration are necessary when attempting to design a space mission, both

system and mission integration. The first, system integration, deals with putting all the

pieces together, essentially the subsystems, into a working vehicle. In our case, this could

be the OTV, the Lander, or the Space Ball. Once all of these are properly integrated, it is

necessary to integrate the set of vehicles into one mission, therefore the term mission

integration. The first step in the process we undertook was to design the subsystems

present in each vehicle. In general each system is designed to minimize its mass

contribution toward the craft‟s total mass. Once the pertinent dimensions are obtained,

the problem of system integration then becomes a much more comprehensive issue.

Several priorities must be considered when integrating a system. The most important

consideration is the center of mass of the set of vehicles. During Lunar Transfer, the

electric thruster must be oriented such that it fires through the center of mass. However, if

an interior component is moved ever so slightly, the center of mass of the vehicle moves

off that line of action. Therefore it is important to maintain a center of mass along that

line of action. Also of importance is thermal control of the mission. All components must

be placed in a position where the thermal control system can cool or heat these

components. An iterative process then ensues to ensure that this happens. Finally, of great

importance in a design study such as this is whether or not this vehicle could actually be

manufactured and assembled in the manner in which it is designed. This is an issue that

requires a more intuitive approach. While we are certainly not well versed in

manufacturing principles, we endeavored to make sure that we could actually construct

the structures involved with this mission.

In the process of integrating the OTV, it became clear early on in the design that it would

be beneficial to put all of the electronics equipment in one electronics module. This was

done so all of the equipment giving off heat could be addressed by one common thermal

control system. After the module was designed, the electronics components such as the



DC converters are then all arranged in the module such that they are supported by the

floor beams but still maintain an axial center of mass relative to the spacecraft as a whole.

From here the heat pipes and radiators are attached, as their positions were determined by

the needs of the electronics module to eliminate heat and eject it into space. At this point,

the frame of the OTV, the electronics module, and the thermal control systems were all in

place as we see in Fig. 5.6-1.

Fig. 5.6-1 OTV Frame.

(Korey LeMond)

As no other component needed to be integrated that was not axisymmetric about the line

of action of the electric thruster, the integration process was simply a puzzle of where

these subsystems could be attached on the vehicle. An example of this principle is the

attachment of solar arrays. The array is comprised of two panels that deploy directly

opposite each other relative to the axis of the OTV. Therefore neither will move the

center of mass off the thruster‟s firing line, and it matters relatively little where they are

attached.



The Lander also required integration as a vehicle system. This process occurred in much

the same way as the OTV. The frame was designed to take the loads applied at launch,

and then the subsystems were attached with the same priorities, aligning the center of

mass along the engine‟s firing line and overall thermal control of the vehicle. The main

difference in the integration of the Lander is the thermal control system. This vehicle was

pressurized to establish a gaseous environment more readily controlled, which in turn

caused extensive problems that will be detailed more in the next few paragraphs under

mission integration.

When considered individually, each of these vehicles, the OTV, Lander, and the two

Space Balls, were all fairly self sufficient entities. This greatly reduces the thought that is

put into the mission integration as a whole. Several problems did exist however. The first

was the fact that power had to be transmitted to the Lander during Lunar Transfer. This

meant that a cable had to run from the OTV to the Lander during that mission phase.

Ordinarily this is not a problem. However, with the additional knowledge that the Lander

was pressurized, creating a connection that will separate when the Lander separates from

the OTV without compromising the integrity of the Lander is a much more difficult

problem. We found that it is possible to use a fastener port that essentially plugs into a

socket on the Lander, thus creating a power line from the power system of the Lander to

that of the OTV. In effect this would look like the OTV was unplugging from an

electrical socket when the vehicles separate, much like it looks when one unplugs a

simple household appliance.

Another issue with mission integration was the fact that the brain of the mission, the

CPU, was located on the Lander. This means that the OTV is effectively „dumb‟ upon

launch. This was solved by inserting a fiberoptic cable into the power feed line, in

essence killing two birds with one stone.



The final problem with integrating the OTV and Lander was the fact that they are not the

same size diameter. The Lander has a diameter over half a meter smaller than the OTV.

This then called for an adapter skirt seen in Fig. 5.6-2.

Fig. 5.6-2 Adapter Skirt.

(Korey LeMond)

This skirt had to be strong enough to take the compression launch loads, while also stiff

enough to take the bending stresses caused by the Lander wanting to move laterally due

to launch loads on top of the OTV. The main problem then became connecting the

Lander and OTV to this skirt. We choose to use a series of explosive pyrobolts to connect

the Lander and OTV to the skirt. A pyrobolt is a simple explosive composed of

ammonium nitrol and TNT in small amounts that will gently push the Lander and skirt

apart.

The final integration problem is then integrating the system of vehicles to the fairing atop

the launch vehicle. This was accomplished using a circular steel ring riveted into the

OTV floor with pyrobolts that was then attached to the fairing wall. The steel ring is then

fired off the floor of the OTV using the pyrobolts when the vehicle tumbles out of the

fairing in low earth orbit. The final mission configuration can be seen with an appropriate

scale, without the shear paneling, in Fig. 5.6-3.



Fig. 5.6-3 Final Mission Configuration.

(Korey LeMond)



Fig. 5.6-4 Final Mission Configuration with dimensions.

(Korey LeMond)

Mission Configuration – 100g Payload – Risk Analysis Section 5.7,


5.7 – Risk Analysis

The probability of success of the mission is computed by combining the probability of

success for each vehicle in the mission. We add a redundant space ball in order to boost

the probability of success for the mission. Table 5.7-1 describes the probability of

success for each vehicle.

Table 5.7-1 Vehicle Success Rate

Vehicle Success Rate

Dnepr Launch Vehicle 94%

Orbital Transfer Vehicle 88%

Lunar Lander 88%

Space Balls (redundant) 98%

Mission Success Rate 72%

In order to satisfy the 90% mission success rate requirement, a contingency launch of a

duplicate mission is planned. This duplicate mission will be launched immediately

following any failure of the first mission.

Table 5.7-2 Mission Success Rate

Number of Missions Success Rate

One 72%

Two 92%

Although the individual mission success rate is not acceptable, planning a contingency

launch artificially increases our overall system reliability to satisfy the mission

requirement. The necessity of a contingency launch is closely connected to total mission

cost – a lower mission success rate allows for a less costly mission. We are taking a

gamble that our first mission will work correctly and win the GLXP purse. If our first

mission fails, it is prudent to fix the problem, build, and launch a duplicate mission to win

the GLXP purse in order to reduce capital loss.

Mission Configuration – 100g Payload – Cost Analysis Section 5.8,


5.8 – Cost Analysis

Table 5.8-1 Total Mission Cost

Expense Cost ($M)

Dnepr Launch Vehicle 4.80

Orbital Transfer Vehicle 8.69

Lunar Lander 4.93

Space Balls 0.10

Overhead 8.58

Total Mission Cost 27.1

$/kg Payload 270.9

Our costing method tends to underestimate the total mission costs. We anticipate actual

mission costs will be higher than our estimate. Table 5.8-2 tabulates net profit if we

consider the GLXP purse money.

Table 5.8-2 GLXP Purse and Relative Cost

Expense Cost ($M)

Total Mission Cost (27.1)

GLXP Purse 22.3

Net Profit (4.8)

We anticipate a net loss in capital in this mission.

Mission Configuration – 10kg Payload Section 6,


6 – Mission Configuration 10kg Payload

Mission Configuration – 10kg Payload – System Overview Section 6.1,

Author: Caitlyn McKay


The 10kg payload system configuration meets the mission requirements of surviving the

lunar night, taking then sending pictures and videos, finding a heritage site and moving

500 meters. The system consists of an Orbital Transfer Vehicle and a Lunar Lander, Fig.

6.1-1. The heritage site we choose is where Apollo 12 landed. The Lunar Lander

descends onto the surface close enough that pictures are able to capture the site. The

Lunar Lander takes 8 minutes of high-definition footage and a picture of the panoramic

view. The Lunar Lander will then move by “hopping” from the landing site to a safe area

at least 500 meters away as we see in Fig. 6.1-2. After more pictures have been taken,

the Lunar Lander will shut down for one lunar night. After the lunar night the Lunar

Lander turns back on and sends a signal to allow us to know that it survived. The

complete mission timeline, after the 365 days to travel to lower lunar orbit, we see in

Table 6.1-1 and the masses of each stage in Table 6.1-2.

Fig. 6.1-1 Overview of 10kg payload mission stack

(Christine Troy)



Fig. 6.1-2 Schematic of the Lunar Lander

hopping 500 meters.

(Joshua Elmshauser)

Table 6.1-1 Mission timeline

Elapsed Time

ddd:hh:mm

Event

-365:00:00 Launch

0:00:00 Lunar Lander reaches LLO and separates from OTV

0:00:04 Lands on lunar surface and starts video taping

0:00:12 Finishes taping and begins transmission of video

0:03:44 Completes video transmission and takes panoramic pictures

0:03:45 Finishes panoramic pictures and begins transmission of pictures

0:03:59 Completes picture transmission and begins hop for locomotion

0:04:01 Locomotion phase complete and begins HD video taping

0:12:01 Begins transmission of HD video and takes panoramic pictures

2:06:24 Ends transmission of HD video and begins transmission of pictures

2:06:36 Ends transmission of pictures and shuts down for lunar night

15:23:24 Turns on and sends signal after lunar night. Footnotes:

Elapsed time is in days:hours:minutes.

The time begins when the Lunar Lander is in lower lunar orbit and is detached from

the OTV.



Table 6.1-2 Mission Masses

Phase Mass in kg

Mass delivered to LEO 583.5

Mass delivered to LLO 228.3

Mass delivered to surface 107.0

Payload delivered 10

We designed to place the 10kg payload on the surface with the least amount of mass.

Typically the greater the mass, the more it cost to put an object on the moon. The

following sections give a more detailed look at our design.

Mission Configuration – 10kg Payload – Launch Vehicle Section 6.2.1,



6.2.1 – Launch Vehicle/Site

We select the Dnepr-1 rocket to place the 10 kilogram payload into a low Earth Orbit.

The configuration for this case only weighs 127 kilograms more compared to the 100

gram payload mission. This mass difference does not exceed the Dnepr‟s capabilities,

thus the same performance standards presented in section 5.2.1 can be applied here.



6.2.2 – Earth Parking Orbit Selection

Atmospheric Drag

We analyze the effects of atmospheric drag for the 10 kg payload in the same manner as

for the 100 g payload in Section 5.2.2. We reach the same conclusions and select a

minimum parking orbit of 400 km.

Parking Orbit Selection

Similar to Section 5.2.2, we choose the lowest altitude parking orbit possible. We select

a 400 km circular parking orbit as it overcomes drag effects and yields the minimum cost.



6.2.3 – Attitude determination in LEO

This mission configuration uses the same attitude determination subsystems and

methodology as outlined in Section 5.2.3 with no changes.

Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.1,

Author: Tim Rebold

6.3 – Lunar Transfer

6.3.1 – Structure

Much of the information presented in this section contains differences between Orbital

Transfer Vehicle (OTV) configurations in the 100g and 10kg payload missions. We

leave out all the similar ideas and material presented for the 100g payload mission. For a

better explanation on various topics discussed in this section please refer back to Section

5.3.1. Refer to Section A-6.3.1 for the details and methodology used to determine our

design for the 10kg payload mission.

Limit Loads

Launch loads or limit loads are obtained from the Dnepr User‟s Manual, and used to

evaluate the capability of our spacecraft design just as in the 100g payload case.

Similarly, we only concern ourselves with the launch loads for sizing and designing our

spacecraft, because we assume they will provide an upper bound limit for our vehicle.

Table 6.3.1-1 shows these loads for the Dnepr launch vehicle. The loads specified are

valid only if the spacecraft meets certain stiffness, or natural frequency requirements.

These requirements are summarized in Table 6.3.1-2 for the Dnepr launch vehicle.

Table 6.3.1-1 Dnepr launch limit loads

Event

Axial

Acceleration

Lateral

Acceleration

1st Stage Burn: Maximum Lateral Acceleration 3.0 ± 0.5 0.5 ± 0.5

2nd Stage Burn: Maximum Longitudinal Acceleration 7.8 ± 0.5 0.2

Footnote: Table based from Dnepr User‟s Guide

Table 6.3.1-2 Dnepr spacecraft stiffness requirements

Thrust (Hz) Lateral (Hz)

20 10

Footnote: Table based from Denpr User‟s Guide


Author: Tim Rebold

Margin of Safety

We apply the same factor of safety (FS) of 1.5 when we conduct our analysis. The yield

margin of safety (MS) which we use in all of our analyses is shown as:

𝑀𝑆 = 𝐴𝑙𝑙𝑜𝑤𝑎𝑏𝑙𝑒 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠)

𝐷𝑒𝑠𝑖𝑔𝑛 𝑦𝑖𝑒𝑙𝑑 𝑙𝑜𝑎𝑑 (𝑜𝑟 𝑠𝑡𝑟𝑒𝑠𝑠) − 1 (6.3.1-1)

Configuration

The configuration and layout for the OTV is based on the same design as the 100g

Payload mission. Changes in this design include different thicknesses of individual

structural members to handle different loads produced from a different sized payload.

The main difference between the two missions comes from the change in payload size.

During launch the payload will produce forces on the OTV. The payload (Lander

Lander) becomes heavier in the 10kg payload mission but has a lower center of mass. As

a result most structural members in the 10kg payload mission will experience loads

different of what they experienced in the 100g payload mission. Below we discuss the

changes to individual components and systems.

Stiffeners

The stiffeners (C-Channels) remain the same length with changes made to their cross

section. Since the Lunar Lander (LL) has a lower center of mass in the 10kg Payload

mission (while being slightly heavier), the C-Channels have less of a structural

requirement. This is because bending moments are reduced from the lower center of

mass, which will raise the critical buckling load of each individual stiffener. Therefore,

the stiffeners can be designed lighter and meet their structural requirements.

Propulsion Module

The propulsion frame remains the same as in the 100g Payload mission. See Section

5.3.1 for details on the configuration and layout of the support frame.


Author: Tim Rebold

Electronics Module

The electronics module (E-MOD) remains the same as in the 100g payload mission. See

Section 5.3.1 for details on the configuration and layout of the support assembly.

Skin

The skin also remains the same as in the 100g payload mission. See Section 5.3.1 for

details on the configuration and layout of the skin

Integration

Integration structure includes the Payload Attach Fitting (PAF) and LL skirt. A complete

description of their geometry and functions can be found in section 5.3.1. Both change

due to the change in payload mass and center of mass. The thicknesses in the skirt walls

are optimized to reduce weight while meeting their structural requirements. The LL skirt

ends up getting heavier in this design because the walls need to be thicker to prevent

buckling. This is a result of a heavier payload mass. The PAF ends up getting lighter in

this design since its walls are thinner. A lighter PAF results from the natural frequencies

of the spacecraft being higher because of a lower center of mass. We use the PAF to

drive these fundamental frequencies high enough to meet the Dnepr stiffness

requirements. We also figure that we can add as much mass to the PAF as necessary

since it is separated from our spacecraft after the launch is finished. The PAF itself is

strong and stiff enough that buckling and yielding are not a concern. Detailed geometry

of the LL skirt and PAF can be seen in Section A-6.3.1.

Sizing

The basic principles that we use to size the OTV in the 100g payload mission are also

used in the 10kg payload mission. See Section A-5.3.1 for complete details.


Author: Tim Rebold

Material Selection

The 10kg orbital transfer vehicle is made of the same materials that the 100 gram OTV is

fabricated from. The floor and shear panels are fabricated from magnesium AZ31, while

all other structural members are made of aluminum 6065. All bolts, rivets, fasteners, and

connectors are made of AIS 1015 low carbon steel.

Finite Element Analysis (FEA)

Information presented in this section contains the differences in the 10kg Payload Orbital

Transfer Vehicle (OTV) design versus the 100g payload OTV design. Refer to Section

5.3.1 for a complete explanation of the analysis. Section A-5.3.1 presents all the details

and methods used to conduct our analysis.

FEA Analysis Results

As for the 100g payload OTV design, we perform a similar finite element analysis (FEA)

for our 10kg payload design. We can perform a similar analysis, because the

configuration of the OTV remains almost identical for both designs. The only substantial

difference from the previous mission is in the payload characteristics. Table 6.3.1-1

shows the structural component masses as a result of these analyses.

Table 6.3.1-1 Component mass totals for OTV design


PAF* 41.36

E-MOD floor beams & overlay 5.25

Shear / Skin Panels 15

Propulsion Support Frame 3.01

Stringers / Stiffeners 12.12

Lander Skirt 14.24

Fasteners (welds, rivets, bolts, adhesives)** 2.12

TOTAL 51.74

Footnotes: Not included in final OTV mass

Estimate

Mission Configuration – 10 kg Payload – Lunar Transfer Section 6.3.2,

Author: Levi Brown


Trajectory

We design the trajectory for the 10 kg payload similar to the 100 g payload. The 10 kg

payload results in slightly different initial conditions as seen in Table 6.3.2-1.

Table 6.3.2-1 10 kg Payload Trajectory Configuration

Parameter Value

Payload Mass (kg) 228.1

Thrust (mN) 104

Mass Flow Rate (mg/s) 5.3

Earth Phase Angle (deg) 83

Moon Phase Angle (deg) 221

Parking Orbit Altitude (km) 400

Capture Orbit Altitude (km) 25

Initial Mass (kg) 585.6

Flight Time (days) 365

This model inherently contains a mismatch in position and velocity at the intersection of

the spiral out and spiral in curves (see Fig. 6.3.2-1). We calculate the propellant mass

required to produce the ΔV mismatch operating the main engine. We add this propellant

mass to the OTV to account for the error in position and velocity. We assume that

performing small maneuvers throughout the trajectory eliminates the fairly large bias in

position and velocity at the intersection point. Table 6.3.2-2 contains these mismatch

values.

Table 6.3.2-2 10 kg Payload Trajectory Configuration

Parameter Mismatch

Position (km) 687.6

Velocity (km/s) 435.1

Propellant Mass (kg) 13.1

Figure 6.3.2-1 illustrates the resultant trajectory.

Mission Configuration – 10 kg Payload – Lunar Transfer Section 6.3.2,

Author: Levi Brown

Fig. 6.3.2-1 10 kg Payload Trajectory.

(Levi Brown)

Lunar Capture Orbit

The lunar capture orbit is the same as the 100g payload case, see section 5.3.2.


Similar to Section 5.3.3, we calculate the propellant necessary to perform a 50 m/s burn

with the main engine. The OTV has larger mass for the 10 kg payload, which results in

more propellant required for the correction as seen in Table 6.3.2-1.


Parameter Value

Isp (s) 1964

mo (kg) 585.6

Propellant for Correction (kg) 1.5

-1 0 1 2 3 4

x 105

-2.5

-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

x 105

xhat

(km)

yhat(k

m)

Spiral Out and In of Spacecraft

Spiral Out from Earth

Spiral In to Moon


Author: Brad Appel


Propulsion Subsystem Overview

Although the OTV sees a 46% increase in payload mass in this mission as compared to

the 100g payload case, this actually has a minimal effect on the Propulsion hardware. The

only differences are the power required and total Xenon propellant required. Table 6.3.3-

1 summarizes the propulsion system parameters for our 10kg mission.

Table 6.3.3-1 Propulsion System Totals


Wet Mass 215 kg

Dry Mass 30 kg


Burn time 365 days

Thrust 104 mN


Mass flow Rate 5.4 mg/s


Author: Brian Erson

6.3.4 – Attitude

Attitude Control

As mentioned in Section 5.3.4, three systems are integrated to perform attitude control of

the OTV.

1. Sensors – The same set of sensors are used for all payloads: a sun sensor and star

sensor, both made by VFCT

2. Reaction Wheels – Due to the increased mass of the OTV, a larger reaction wheel

must be used to maintain attitude stability throughout the trans lunar phase.

VFCT manufactures the VF MR 10.0, a wheel capable of creating adequate torque

for the system.

3. Thrusters – The 𝐻2𝑂2 thruster hardware needed is the same as the 100g case.


complete system accomplishing the mission of controlling the OTV during the translunar

phase. Below, Table 6.3.4-1 provides an overview of our ACS specifications for the 100g

payload.

Table 6.3.4-1 OTV ACS Budget for 10kg payload

Device Mass (kg) Cost ($) Power Required(W)

VF STC 1 (star sensor) 6.4 133,333 20.4

VF SNS (sun sensor) 0.7 133,333 5

VF MR 10.0 20 133,333 120

𝐻2𝑂2 thruster 0.36 1,500 --

𝐻2𝑂2 Propellant 1.26 100 --

Inert Mass 3.02 1,000 --

Totals 31.74 403,000 145.4


Author: Brian Erson


As mentioned in Section 5.3.4, 𝐻2𝑂2 is an adequate propellant to fulfill all mission

requirements.


The analysis for the 10kg payload is identical to the analysis for the 100g payload

detailed in Section 5.3.4.





For the 10kg payload we will be using the same systems as for the 100g payload. For

more information refer to Section 5.3.5 and Section 5.4.5.

Communication Link Budget

Please see Section 5.3.5 for the communication link budget during the lunar transfer. The

communication requirements and equipment are unchanged with the change in vehicle

and payload mass.


For the 10kg payload the ground station usage is the same as the 100g payload and can be

found in Section 5.3.5.

Communication Antenna Pivot

The patch antenna pivot discussed in Section 5.3.5 will be used again for the 10kg

payload case, and will be essentially unchanged.





Similar to the 100g payload case, we thermally control the electronics, the xenon tank,

and the electric thruster for the 10kg payload OTV. The electronic components put out a

total of 283W of heat. To cool these components, we employ the same thermal control

system in the 100g case. In Table 6.3.6-1, we see the mass and volume breakdown of the

thermal control components for the electronics.


For the 10kg payload case, we take advantage of a simple wire heater and multi-layer

insulation to keep the xenon within its storage temperatures.


The electric thruster generates approximately 849W of heat that we need to dissipate to

keep the thruster under its maximum operating temperature of 473K (200ºC). We see the

dimensions of the OTV‟s thermal control system in Table 6.3.6-1.

Table 6.3.6-1 10kg Payload OTV Thermal Control Dimensions



Heat Pipes 2.53 ID = 3.83cm; OD = 4.13cm;

Length = 5m

Radiators 1.576 Total Cross-Sectional Area =

1.168m2

(8 fins @ 0.146m2 each)





6.3.7 – Power

Similar to the 100g payload case, the OTV for the 10kg case is powered by two circular

Ultraflex solar arrays and a secondary (rechargeable) lithium-ion battery. Due to the

increased power requirements, however, the dimensions of both the solar arrays and the

battery change. We see a CATIA computer model of one of the partially deployed solar

arrays in Fig. 6.3.7-1.

Fig. 6.3.7-1: Top and side views of OTV Ultra-flex solar array CATIA model; partially deployed.

(Ian Meginnis)

Power Budget

We see the adjusted power budget for the OTV during the Lunar Transfer phase in Table

6.3.7-1. The largest contributing factor to the OTV budget still arises from the

propulsion group‟s electric thruster. This device consumes over 75% of the entire OTV

power budget. To adequately meet all of the power needs, we increase the total power

production by 5%.



Table 6.3.7-1 10kg Payload OTV Power Budget by Group

Group Power Units




Power 120 Watts

Lunar Lander

(during Lunar Transfer) 105 Watts

TOTAL 2534 Watts

Solar Array Sizing

With the power budget, we determine the sizes of the solar arrays. Since the solar arrays

for the 10kg payload case are merely scaled up, we employ the sizing methods from the

100g payload case. We see the mass, stowage volume, and deployed area for the two

solar arrays in Table 6.3.7-2.

Table 6.3.7-2 10kg Payload OTV Solar Array Dimensions (total)


Mass 16.89 kg

Stowage Volume 0.0858 m3


The manner in which the solar arrays are stored and deployed is identical to the 100g

payload case. To track the sun during the Lunar Transfer phase, we include two motors

and a sun sensor.

Similar to the 100g payload case, one of the largest driving factors in determining the

OTV‟s mission configuration for the 10kg case is the high cost of the solar arrays. The

solar arrays cost a total of $2.53 million.



Battery Sizing

The lithium-ion battery cells for the 100g payload case are also employed for the 10kg

case. We see the new dimensions of the battery in Table 6.3.7-3. Although the OTV‟s

battery still comprises of the same types of lithium-ion cells, we have 13 cells instead of

ten cells. This results from the increased battery energy capacity.

Table 6.3.7-3 10kg Payload OTV Battery Dimensions



Volume 0.0056 m3


The components for the OTV‟s power system for the 100g payload case are also used for

the 10kg case. We reference Fig. 5.3.7-2 in Section 5.3.7 for a diagram of the OTV

power system components. Table 6.3.7-4 shows the dimension of the 10kg payload

power system components.

Table 6.3.7-4 10kg Payload OTV Power Subsystem Dimensions



Battery 15.93 0.0056

PCDU 11.49 0.0203


Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1,




CAD

We see in the following figures a CAD model created in the program CATIA. This model

gives us a visual representation of what our Lander looks like, and gives us a general idea

of where certain systems are located. The Lander assembly is designed with a color

coding system in mind to easily identify certain subsystems.

7) Red – Structures

8) Orange – Attitude

9) Yellow – Power

10) Green – Communication

11) Blue – Propulsion



Fig. 6.4.1-1 External View of the Lunar Lander with legs deployed.




Fig. 6.4.1-2 Close up of the Lunar Lander’s internal components with legs stowed.




Fig. 6.4.1-3 View of the Lunar Lander’s engine systems.




Fig. 6.4.1-4 Visualization of the 10kg Lander on the Moon.



Author: Ryan Nelson

Structure

The most important consideration in designing our structure of the Lunar Lander for the

10kg payload is the mass of the frame. We identify the size and mass of the Lunar

Lander at the first lunar touchdown as driving factors for the total mass of the Lunar

Lander frame structure.

The two additional engines needed to do the hop for the 10kg payload case increase the

size of the Lunar Lander frame. The height of the chambers for the hop engines is larger

than the height of the chamber for the main engine. The height of the Lunar Lander is

increased from the 100g payload case partially because of the large chamber of the hop

engines. However, the main reason for designing a Lunar Lander frame height of 1.1m

(as depicted in Table 6.4.1-1) is because the line of fire for the hopper thrusters needs to

be through the center of mass at initial lunar touchdown. The height increase of the

Lunar Lander to 1.1 meters allows for the placement of several Lunar Lander

subcomponents higher up in the frame. This increases the height of the center of mass for

the Lunar Lander. Since the center of mass is higher, the angle at which these thrusters

fire is smaller which results in the need for propellant. This results in the reduction of the

entire Lunar Lander mass including the frame. The increase of the Lunar Lander height

brings the total volume of the Lunar Lander to 1.15 m3.

Table 6.4.1-1 Basic Frame Dimensions for 10kg payload Lunar Lander


Height 1.1 meters

Bottom Diameter 1.3 meters

Top Diameter 1.0 meters

Length of Legs 0.606 meters

The basic shape of the Lunar Lander for the 10kg payload case is a conic frustum.

Figure 6.4.1-1 illustrates the main frame components for this Lunar Lander. The frame

floor consists of a circular outer ring and circular inner ring connected by four rectangular


Author: Ryan Nelson

floor supports. Another circular ring makes up the top of the Lunar Lander frame. This

top circular ring connects to the frame floor by four hollow circular side support beams.

These side support beams are able to support the various loads subjected to the Lunar

Lander throughout the mission while maintaining a low thickness. A low thickness

enables the storage of the four Lunar Lander legs within these side supports during Earth

launch and Lunar Transfer. A 0.5mm magnesium skin around the entire Lunar Lander

frame protects against micrometeorites and provides thermal protection. The lower

density of magnesium than aluminum makes magnesium more attractive for the Lunar

Lander skin.

Fig 6.4.1-1 Schematic of Basic Lunar Lander Frame Components. (Ryan Nelson)

As depicted in Fig. 6.4.1-1, the addition of the two engines needed to make the hop

results in some extra structural mass to connect these engines. The mass needed to

integrate the two engines to the main frame of the Lunar Lander is about 5 kg. The

addition of this extra structural mass for engine integration combined with the total mass

of the Lunar Lander at initial touchdown brings the frame mass to 19.97 kg.



Material Selection

The 10kg Lunar Lander is made of the same materials that the 100 gram Lunar Lander is

fabricated from. The floor and shear panels are fabricated from magnesium AZ31, while

all other structural members are made of aluminum 6065. All bolts, rivets, fasteners, and

connectors are made of AIS 1015 low carbon steel.




Overview

In Section 5.4.2, we described the descent profile of the Lunar Lander for the 100g

payload. The descent profile of the 10 kg payload is very similar. Differences arise in

the thrust of the descent main engine, burn times, amount of propellant used, and perilune

altitude where final descent begins. In this case, descent begins at an altitude of 13.8 km.

Tables 6.4.2-1 and 6.4.2-2 highlight the additional differences for the 10 kg payload.

Table 6.4.2-1 Final Descent Action Breakdown


Radial Burn 000 - 177 1650

Rotation 177 - 183 165

Vertical 183 - 199 1650

Coast to Land 199 - 210 180



Mass (kg)

De-Circularize

0.5

Radial Burn 000 - 177 97.8

Rotation 177 - 183 0.3

Vertical 183 - 199 8.5

Coast to Land 199 - 210 0.9

Unusable

3.5

Contingency

1.0

Total

112.5

Footnote:

An additional 8.7 kg of propellant is needed for

the hop, bringing the total propellant mass in

LPO to 121.2 kg.



Landing Location

We chose to land at the same landing site as the 100g payload case. Please refer to

Section 5.4.2.




Overview

The Lunar Lander main engine we designed for the 10kg payload is of the same type and

configuration as the engine described in Section 5.4.3 for the 100 g payload case. Fig.

6.4.3-1 shows the engine dimensions resulting from higher thrust requirements and

includes the 100g payload Lunar Lander engine for a size reference.

Fig. 6.4.3-1 Main Engine Dimensions.

(Thaddaeus Halsmer)


When we scale the main engine for the 10kg payload case, only the thrust and burn time

change to the specifications given in Table 6.4.3-1. Isp, chamber pressure and nozzle area

ratio are the same as given in Section 5.4.3.



Table 6.4.3-1 Hybrid Engine Performance Specifications


Thrust Max / Min 1650 / 165 [N]

Burn Time 190.4 [s]


The propellant feed system operates in the same manner we discussed in Section 5.4.3,

although three isolation solenoid valves were made necessary by adding the two

locomotion engines that are discussed in Section 6.5.1. The isolation solenoid valves

allow any engine to be purged and fired individually using the common feed system and

oxidizer flow control valve. The fluid system schematic given in Fig. 6.4.3-2 shows the

integration of the additional valves and engines.

Fig. 6.4.3-2 Lander Propulsion Fluid Schematic

(Thaddaeus Halsmer)



Table 6.4.3-2 Additional Fluid control components not in Table 5.4.3-2

Item Description

SV03 Locomotion engine #1 ISO Solenoid Valve

SV04 Main Engine ISO Valve

SV05 Locomotion engine #2 ISO Solenoid Valve


Author: Saad Tanvir


Adding the three solenoid valves to the propellant feed system increases the maximum

power requirement to 150 watts.

H2O2 Storage

Employing the same technique as in Section 5.4.3, we determine the amount of H2O2

decomposed over the period of one year. The hydrogen peroxide is stored in a sealed

spherical Aluminum 1060 tank. The amount of H2O2 decomposed during the mission is

0.121 kg. As compared to the 121 kg of total propellant carried, this value is negligible.



the H2O2 tank.




Volume 0.0867 m3

Thickness 1.26 mm

Diameter 0.550 m

Mass 3.35 kg


the mission (0.0834 m3: including the usable propellant volume for: main engine +

hopper engine, boil-off volume and the volume of unusable propellant) and a 4% ullage

volume (0.00333 m3).



the material for the He tanks (Humble). Table 6.4.3-4 shows the detailed specifications

of the pressurant tank as well as the mass of the Helium gas required of the whole landing

mission.


Author: Saad Tanvir




Volume 0.0397 m3

Thickness 6.70 mm

Diameter 0.41 m

Mass of the Tank 1.34 kg

Mass of the Helium 1.37 kg


The inert mass of the propulsion system is obtained as described in Section 5.4.3. The

inert masses for the additional locomotion engines are largely dependent on the amount

of propellant we use during the locomotion (hop) and the average thrust the engine

produces during locomotion. Using the propellant masses and the average thrust obtained

from mission operations we size the additional nozzles, chambers and feed system for the

two locomotion engines. The oxidizer (H2O2) is obtained from the same tank that

provides H2O2 to the main Lunar Lander engine. The total propulsion system inert mass

for the hybrid landing engine plus the two additional locomotion engines equals 45.67 kg.

Table 6.4.3-5 gives a detailed breakdown of the propulsion system inert masses.


Author: Saad Tanvir



Main Lander Engine




Nozzle 15.74 kg


Injector 0.99 kg


Feed System 0.99 kg


Valves 6.94 kg

Hopper Engines

Nozzles 6.30 kg

Combustion Chambers 0.29 kg

Injectors 0.09 kg


Feed Systems 0.09 kg


Valves 1.10 kg


Thermal Analysis

As in Section 5.4.3, similar thermal analysis is performed for this case. We find that to

keep the H2O2 at 283 K temperature during the lunar transfer phase we require a constant

power supply of 37 W.

During lunar descent, the temperature drop in the hydrogen peroxide tank is 5.8 K (from

283 K to 277.2 K). Similar temperature drop from the valves and the feed lines are 2.2K

and 2.0K respectively. Because the drop in temperature of these components during lunar

descent is small, its affect on engine performance is negligible. Hence, there is no power

required for thermal control on the propulsion system during lunar descent.


Author: Josh Lukasak


The attitude control system for our10kg payload case is identical to the system we use on

the lunar descent vehicle in the 100 gram payload case. For both payload cases the

attitude control system is used to counteract the moment that the lunar descent engine

creates. The main engine creates a torque on the spacecraft when the line of action of the

thrust does not pass through the center of mass of the descent module. Therefore we use

the same attitude system for both cases due to the similarities in the amount of thrust

generated by the lunar descent main engine. We also use the same attitude detection

devices from the previous case for the 10kg payload case. The sun sensors and star

sensors are located on the lunar descent vehicle for the duration of the mission and are

used for both the orbital transfer phase and the lunar descent phase. The configuration of

the thrusters on the lunar descent vehicle is found in the attitude control section for the

100 gram payload. The amount of attitude propellant is the only portion of the attitude

control system that changes between the 100 gram and 10kg payload cases. The total

mass for the attitude control system masses can be seen in Table 6.4.4-1.

Table 6.4.4-1 Attitude System Masses

Attitude Component Mass (kg)

De-Orbit Burn Attitude Propellant 0.5

Descent Attitude Propellant 5.13

Inert Attitude Control 10.18

Total 15.71





The communication equipment we are using on the Lander for the 10kg payload mirrors

the 100g payload very closely, which can be found in Section 5.4.5. However, the 10 kg

payload differs from the 100g payload in one place. While we need two communication

systems for the 100g payload we now only need one system because the Space Balls are

not used. Table 6.4.5-1 lists the total mass, power usage and purchase price for the

system.

Table 6.4.5-1 Communication Equipment Totals

Mass (kg) Power Usage (W) Purchase Price (2009 $)

2.17 54.53 295,167

The communication systems consist of:

1) Two antennae to communicate with Earth.

2) Two antennae mount pivots.

3) A transmitter to transmit to Earth.

4) A receiver to receive signals from Earth

5) A video camcorder to record video and take pictures

6) A Computer board to control these systems.


Please refer to Section 5.4.6 for details regarding the communications power

requirements for the Lunar Lander/Hopper.


Please see Section 5.4.5 for the communication timeline and GLXP requirements for the

10kg payload case, as they have gone unchanged and the way that we will achieve them

is unchanged as well.




Lunar Day

The thermal control for the 10kg payload case is similar to the thermal control for the

100g payload case. We describe the main system in the 100g payload case, Section 5.4.6.

This section describes the differences between the two payload cases. Fig. 6.4.6-1 is the

diagram of the day thermal control system.

Fig. 6.4.6-1 Schematic of Passive Thermal Cooling System.

(Josh Elmshaeuser)

We use the multi-layer insulation (MLI) blanket on the entire outside of the Lunar

Lander, like the 100g payload case. The difference between the two cases is the

locomotion mode. The MLI blanket needs to cover the hopper engines. Also, the hopper



engines will need a resistance heater. The total system needs 9.5 Watts during Lunar

Transfer.

The other sections of the Lunar Lander thermal control will be the same. The length of

the heat pipe remains at five feet, and the communication plate remains the same size. All

the other thermal control devices scale based on the size of the Lunar Lander. A

breakdown of the subsystem masses are in Table 6.4.6-1. The total mass of the system is

9.56 kilograms. The total mass was calculated in the MATLAB program

LanderThermalControl.m.

Table 6.4.6-1 Day Thermal Control of Lunar Lander


MLI Blanket 2.38

Aluminum Plate 1.40

Heat Pipe 2.51

Radiators 2.80

Ammonia 0.0215

Heaters 0.45

Total 9.56

Surviving the Lunar Night

We will encounter the same environmental conditions for the 10 kilogram payload

configuration as with the 100 gram payload. The most pronounced difference is the

slightly greater surface area of the protected enclosure so that the heat loss is

approximately 11.5 Watts instead of 10 Watts

We are using the same heating and shutdown system as described in Section 5.4.6. The

only difference is the amount of hydrazine is increased by about 5% as is the tank volume

to compensate for the increased heat loss.

Mission Configuration – 10 kg Payload – Lunar Descent Section 6.4.7,

Author: Adham Fakhry

6.4.7 – Power

Battery

We size the battery for the Lunar Lander based on power requirements from the Lander

engine, communication and attitude systems. The LL engine, the communication gear

and the attitude systems all depend on the battery during the landing phase.

The LL‟s engine requires power to operate the various valves and pumps to make the

engine function, which total 148 Watts and will be operating for 450 seconds, the first

250 seconds is for the landing phase, then the engine is ignited again for 200 seconds

after landing and this is the hopping phase, where the LL travels 500 meters and then

lands again.

The communication gear uses a peak total power of 54.53 Watts, to ensure that we can

maintain constant communication between the Earth and the LL throughout the entire

landing phase. We provide enough power for the communication gear to transmit for 30


Watts to power our system and requires 15 minutes of peak power to make sure the LL

arrives at its destination. Table 6.4.7-1 highlights the total power requirement for all

systems and their individual components on the 100 g Lander.

Footnote:

Complete description of each device can be found in

Section 5.4.2

For a complete background on each device within each system, please refer to Tables A-

6.4.7-1 through Tables A-6.4.7-3 in Section A-6.4.7.


Item Power Units



Attitude 25.4 Watts

Total 227.93 Watts



We purchase the batteries from the company Yardney Technical Products, Inc., who

provide readymade lithium ion batteries that are used in space vehicles.

We select a 21 Ampere-hour battery, but the battery only requires 16.95 Ah. We decided

that as a safety precaution the battery should have a little bit extra power, in case of any

problems that might occur after separating from the OTV. Refer to Section A-5.4.7 to see

how we calculated the capacity of the battery in Ah.

In order to avoid loss of charge when not in use, which happens with all lithium ion

batteries, our battery on the LL will be completely drained during integration with the

entire spacecraft. The team will send a command that will allow the battery to charge

when it approaches the moon, using the OTV solar cells before separation. We compute

the battery size using lander10kg.m.

Solar Array Sizing


week long lunar day. The peak power required is 54.53 W to run all of the

communication equipment, refer to Section A-6.4.7 to see breakdown of power

requirements for each device within the communication equipment.

The design for the solar cells for the 10 kg case is the same as the 100 g case, since the

power requirements decreases by 6 watts and the solar arrays are designed to provide

extra power, we decided to keep solar arrays same size as the 100g LL. Refer to Section

5.4.7 for a detailed analysis of the solar cells. Figure 6.4.7-1 displays the power

consumed and the power generated by the cells through a lunar day.



Fig. 6.4.7-1 Maximum Potential Power of solar cells during a lunar

day. The communication gear has enough power to meet its peak

power requirements on the 1.5st day of the lunar day. At maximum

power, the solar cells provide 74.5% more power than required at

day 7 of the lunar day.

(Adham Fakhry)

The red line represents the power used by the communication gear, the blue line

represents the power produced by the cells. The solar cells provide more than enough

power required by the communication gear. Also, the excess power ensures the dust

accumulated on the lunar surface or impacts from meteorites, will not severely affect the

power collected by the cells as they provide extra power. We sized the solar cells to meet

the communication system‟s power requirements during data transmission from the Earth

to the Moon.

The solar cells are provided by Able Engineering, who will be supplying solar cells for

the Orbital Transfer Vehicle and the Lander. The Lander solar cells will cost $250,000

and the solar cells will also weigh a maximum of 2 kilograms. Solar array sizing is



automated in lander.m. Refer to Section A-5.4.7 for breakdown of costs and the

constants used to size the solar cells.

DC-DC Converters

The DC-DC converters for the 10kg case depend on the number of component within the

system and the voltage they each require to operate, the DC-DC converters change the

voltage from the battery and the solar cells to a higher voltage for each individual device

will require. Tables A-6.4.7-1, A-6.4.7-2 and A-6.4.7-3 in Section A-6.4.7 refer to the

various components that require a DC-DC converter. We choose DC-DC converters for

each device by using the data in Tables A-5.4.7-6 through A-5.4.7-10, which list all the

DC-DC converters we are considering to use. Table 6.4.7-2 below displays the DC-DC

converter system for the 10 kg payload.



Mass 0.815 kg

Dimensions 0.033 x 0.033 x 0.033 m


Cost $ 57,500 - Footnote:

Cost is in 2009 US dollars and is the cost for all the units.





A power conditioning and distribution unit (PCDU) is needed to safely manage the power

distribution throughout the LL during mission. Please refer to Section 5.4.7 for the

complete analysis of why we need a PCDU, what the unit is used for and its

specifications.

Mission Configuration – 10kg Payload – Locomotion Section 6.5.1,

Author: Alex Whiteman

6.5 – Locomotion

6.5.1 – Trajectory

The requirements for the 10kg payload locomotion phase are to travel 500m on or over

the surface of the Moon using a propulsion system on the Lunar Lander to perform a

“hop”. The Lunar Lander begins and ends the trajectory in an upright position and uses

the attitude control system to change orientation during flight. Throttling of the

propulsion system will be used throughout the flight to make the trajectory more efficient

and ensure that the Lunar Lander touches down safely.

In order to keep the overall 90% mission probability of success, the main lunar descent

engine cannot perform the hop. Instead, one hybrid thruster orientated at 32.3° from the

Lunar Lander‟s axis of symmetry provides the necessary thrust to perform the hop. This

thruster is one of a pair needed to have the necessary redundancy in the propulsion

system. This thruster configuration has a significant effect on the trajectory as the Lunar

Lander will need near zero horizontal velocity in order to land. In order to achieve this,

the Lunar Lander will have a horizontal velocity in a direction opposite that of the thrust

during the last part of the hop. This increases the required propellant and the time of

flight compared to the case where a vertical thruster provides thrust for the hop.

Fig. 6.5.1-1 shows the hop trajectory as well as the orientation of the Lunar Lander at

various points in the trajectory. Table 6.5.1-1 provides the results for this trajectory.



Fig. 6.5.1-1 Hop trajectory with orientation visuals at various points during flight.

(Alex Whiteman)

Table 6.5.1-1 Hop Trajectory Results


Time of Flight 134.5 sec

Propellant Used 8.6 kg

Average Thrust 197 N

-1000 -900 -800 -700 -600 -500 -400 -300 -200 -100 0 100-100

-50

0

50

100

150

200

range (m)

altit

ude

(m)

Moon

Hop Trajectory

Liftoff Touchdown




Overview

We use hybrid engines to complete the locomotion requirement of 500 meters. The

locomotion engines use the same chemical propellants and share a common H2O2 feed

system with the Lunar Lander main engine. However, the locomotion engine

configuration is changed to a simple axial flow engine with a single circular port. While

the main engine and locomotion engines share a common oxidizer feed system, they can

each be fired independently of each other. Fig. 6.5.2-1 shows the configuration of the

locomotion engine along with its dimensions.

Fig. 6.5.2-1 Locomotion engine configuration and dimensions

(Thaddaeus Halsmer)


We designed the locomotion engine to provide a median thrust to weight ratio of 1 during

the locomotion maneuver. Like the main engine the locomotion engine can be throttled to

10% of maximum thrust by controlling the H2O2 mass flow rate. Because of the fuel

grain port configuration change and reduced burn time, relative to the main engine, the Isp

of the locomotion engine is higher than the that of the main engine. Table 6.5.2-1 shows

the primary performance specifications of the locomotion engine.



Table 6.5.2-1 Hybrid Locomotion Engine Performance Specifications


Thrust (average) 192 [N]

Isp, vacuum average 330 [s]

Chamber Pressure Max / Min 2.1 / 0.21 [MPa]

Nozzle Area ratio Ae/At 100 --

Burn Time 134.5 [s]


The locomotion engines use the same oxidizer feed system used for the main engine

which was described in Section 6.4.3.


The power requirements during the locomotion maneuver remain unchanged from those

described in Section 6.4.3.



6.5.3 – Attitude

The attitude control thrusters rotate the lunar descent craft and orient it in a fashion that

will allow for optimal hop thruster firing. The attitude sensing and control devices are the

same as for the lunar descent phase. These specifics can be seen in Section 6.4.5. What is

important to note is that the hydrogen peroxide thrusters will only require 0.4 kg of

propellant for the locomotion phase of this payload case. This attitude control propellant

is used not only for counteracting hopper engine torque but also to orient and rotate the

descent vehicle to apply the hopper engine thrust in the proper direction. Exact

specifications on how the attitude control thrusters are used in the locomotion trajectory

can be seen in Section 6.5.2.

Mission Configuration – 10kg Payload – Integration Section 6.6,



Much of the integration process for the 10kg payload case was exactly the same as for the

100 gram payload case. In this case, the only differences lie in the fact that there is no

longer a Space Ball, and two extra engines had to be integrated into the OTV electronics

module. The Lunar Lander has three engines for reliability purposes that have to be

integrated onto the vehicle. We do this by moving the electronics components further

away from the axis of symmetry of the OTV to clear room for the extra nozzles.

Once this arrangement is completed, the entire set of vehicles are integrated together in

the same manner as the 100 gram case and the mission is ready to launch.

Fig. 6.6-1 Final Mission Configuration.

(Korey LeMond)

Mission Configuration – 10kg Payload – Integration Section 6.6,


Fig. 6.6-2 Final Mission Configuration with dimensions.

(Korey LeMond)

Mission Configuration – 10kg Payload – Risk Analysis Section 6.7,




success for each vehicle in the mission. We add a redundant hybrid engine to the hop

phase in order to boost the probability of success for the mission. A table describing the

probability of success for each vehicle is shown below:



Dnepr Launch Vehicle 94%


Lunar Lander 88%

Hop (redundant) 99%


In order to satisfy the 90% mission success rate requirement, a contingency launch of a

duplicate mission is planned. This duplicate mission will be launched immediately

following any failure of the first mission.

Table 6.7-2 Mission Success Rate

Number of Missions Success Rate

One 72%

Two 92%

Although the individual mission success rate is not acceptable, planning a contingency

launch artificially increases our overall system reliability to satisfy the mission

requirement. The necessity of a contingency launch is closely connected to total mission

cost – a lower mission success rate allows for a less costly mission. We are taking a

gamble that our first mission will work correctly and win the GLXP purse. If our first

mission fails, it is prudent to fix the problem, build, and launch a duplicate mission to win

the GLXP purse in order to reduce capital loss.

Mission Configuration – 10k Payload – Cost Analysis Section 6.8,




Expense Cost ($M)

Dnepr Launch Vehicle 5.77


Lunar Lander / Hopper 5.71

Overhead 8.58


$/kg Payload 2.96

Our costing method tends to underestimate the total mission costs. We anticipate actual

mission costs will be higher than our estimate. Table 6.8-2 tabulates net profit if we

consider the GLXP purse money.

Table 6.8-2 GLXP Purse and Relative Cost

Expense Cost ($M)

Total Mission Cost (29.6)

GLXP Purse 22.3

Net Profit (7.3)

We anticipate a net loss in capital in this mission.

Mission Configuration – Large Payload Section 7,


7 – Mission Configuration Large Payload

Mission Configuration – Large Payload – System Overview Section 7.1,



We now modify our Google Lunar X PRIZE (GLXP) mission architecture to deliver an

arbitrarily large payload to the lunar surface. Our goal is to resupply an existing lunar

base. The requirements of this mission are to travel to the moon, descend to within

100 m of the surface, and then travel 500 m over the lunar surface in at least one minute.

For the large payload mission configuration, we launch aboard the Falcon 9 launch

vehicle. Our Orbit Transfer Vehicle (OTV) and Lunar Lander are based on the designs

for the GLXP size payloads, modified to accommodate the larger mass we now carry to

the Moon. Figure 7.1-1 shows an overview of our vehicle configuration for the large

payload case.

Fig. 7.1-1 Overview of mission vehicle stack for large payload

(Christine Troy)



Since our requirements for this mission configuration do not stipulate that we must land

on the lunar surface before beginning our locomotion phase we incorporate a hover phase

to achieve our required locomotion. The Lunar Lander descends to 100 m above the

lunar surface, comes to rest, and then travels laterally 500 m. Figure 7.1-2 shows a view

of our Lunar Lander completing its locomotion phase.

Fig. 7.1-2 Lunar Lander completing locomotion for large payload mission configuration.

(Christine Troy)

The general timeline of our large payload mission is given in Table 7.1-1. The elapsed

times in this table are referenced from the time of arrival into low lunar orbit. It takes the

combined OTV and Lunar Lander stack one year to travel from low earth orbit to low

lunar orbit. We transfer from the Earth to the Moon using an electric propulsion system

and spiral transfer orbit. Descent to near the lunar surface and hover locomotion are

accomplished with our hybrid descent engine. We land in Mare Cognitum, near the

landing site of Apollo 12.



Table 7.1-1 Mission Timeline for large payload configuration

Elapsed Time

(ddd:hh:mm) Event

-365:00:00 Launch

0:00:00 Arrive in Low Lunar Orbit/Separate from OTV

0:00:01 Transfer to Lunar Descent Transfer Orbit

0:00:56 Begin Final Lunar Descent burn

0:00:59 Come to rest 100 m above surface/begin hover locomotion

0:01:00 Touch down on lunar surface

Footnote: Elapsed Time given in days, hours, and minutes

Table 7.1-2 summarizes our key mass deliverables at important waypoints along our

mission timeline. The total mission cost and cost per kilogram of payload are also

included.

Table 7.1-2 Large payload system masses and cost

Parameter Mass Units

Injected Mass to Low Earth Orbit 9953 kg

Injected Mass to Low Lunar Orbit 4545 kg

Mass on Lunar Surface 2325 kg

Payload Delivered to Lunar Surface 1743 kg

Total Mission Cost 222.6 Million $

Cost per kg of Payload to Moon 128,000 $/kg

Mission Configuration – Large Payload – Launch Vehicle Section 7.2.1,



7.2.1 – Launch Vehicle/Site

We select the Falcon 9 rocket to place the large payload into a low Earth orbit (LEO).

The Falcon 9 is a new reusable launch vehicle designed by Space Exploration

Technologies (SpaceX). Although the Falcon 9 has not completed any missions to date,

successful tests on the payload separation mechanism and engine configuration have been

accomplished. Additionally SpaceX has taken extra steps in the design of the Falcon 9 to

increase mission reliability. Specifically the Falcon 9 has been designed to a 1.4 factor of

safety making it safe for human flight. The first and second stages also have significantly

higher margins of safety because they are designed to be reusable. The Falcon 9 is

propelled by ten Merlin 1C liquid propellant rocket engines (nine for the first stage, one

for the second stage). SpaceX has limited the number of stages to two to minimize

separation events and possible failures.

Fig 7.2.1-1. Falcon 9 Launch Profile.

(Falcon 9 Launch Vehicle Payload User’s Guide)



Table 7.2.1-1 Characteristics of the Falcon 9 Launch Vehicle


Height 55 m

Diameter 3.66 m

Lift-off Mass

Mass to 400km orbit

333,400

9,953

kg

kg

Number of Stages

Orbit Inclination

Cost per kilogram to LEO

2

28.5

3,600

--

degrees

$/kg

The Falcon 9 will launch from Space Launch Complex 40 (SLC-40) on Cape Canaveral

Air Force Station (CCAFS) in Florida, USA. Cape Canaveral is located adjacent to the

Atlantic Ocean; which ensures safe stage drops. With 55 launches to date, SLC-40 is also

the former home of the Titan IV heavy lift rockets. Launches from Cape Canaveral are

capable of having an initial orbital inclination between 28° and 57°. The launch site is

located 28° north of the equator making it a good spot to maximize launch vehicle

performance

Fig. 7.2.1-2 Map of Launch Site.

(Google Maps)



7.2.2 – Earth Parking Orbit

Atmospheric Drag

We analyze the effects of atmospheric drag for the large payload in the same manner as

for the 100g payload in Section 5.2.2. We reach the same conclusions and select a

parking orbit of 400 km.

Parking Orbit Selection

Similar to Section 5.2.2, we choose the lowest altitude parking orbit possible. We select

a 400 km circular parking orbit as it overcomes drag effects and yields the minimum cost.


Author: Kris Ezra

7.2.3 – Attitude determination in LEO

This mission configuration uses the same attitude determination subsystems and

methodology as outlined in Section 5.2.3 with no changes.

Mission Configuration – Large Payload – Lunar Transfer Section 7.3.1,

Author: Tim Rebold

7.3 – Lunar Transfer

7.3.1 – Structure

Much of the information presented in this section contains differences between Orbital

Transfer Vehicle (OTV) configurations in the 100g and Large Payload missions. For a

better explanation on various topics discussed in this section please refer back to Section

5.3.1. Refer to Section A-5.3.1 for the details and methodology used to determine our

design for the large payload mission.

Limit Loads

We obtain launch loads or limit loads from the Dnepr User‟s Manual, and use them to

evaluate the capability of our spacecraft design just as in the 100g payload case.

Similarly, we only concern ourselves with the launch loads for sizing and designing our

spacecraft, because we assume they will provide an upper bound limit for our vehicle.

See Section 5.3.1 for details on how we select these loads.

Margin of Safety

We apply the same factor of safety (FS) of 1.5 when we conduct our analysis. See

Section 5.3.1 for details on how we select this factor of safety.

Configuration

The configuration and layout for the OTV is based on the same design for the 100g

payload mission. Changes in the diameter and height of the OTV are made to increase

volume. We increase the height and diameter of the OTV to 2.58 m and 3.6 m,

respectively. We increase the volume of the OTV, so we can fit more equipment and

larger components inside. The propulsion system is much larger in this mission along

with many other mission components. Changes to individual structural members include

different cross section thicknesses and dimensions to handle larger loads produced from a

heavier payload. These changes are illustrated in further detail in Section A-7.3.1. The


Author: Tim Rebold

main difference between the two missions comes from this change in payload. During

launch the payload produces forces on the OTV. The payload or Lunar Lander (LL) is

significantly heavier in the large payload mission. As a result the structural members

experience stronger loads from what they experienced in the 100g payload mission. We

can scale many of the components in this design because of the similarities in OTV

configuration. Below we discuss the changes to individual components and systems.

Stiffeners

The stiffeners in this mission have a larger cross section to deal with larger bending

moments and compressive loads. Unlike the previous two missions, the size of the

stiffener cross section is determined from the bending loads transferred through them.

Buckling was the limiting failure mode in the previous designs, and determined the

needed cross section properties. See Section 5.3.1 for more details on the configuration

and layout of the support assembly.

Propulsion Module

We make the members of the propulsion frame stronger and stiffer to support a heavier

propulsion system. The cross section dimensions will again increase to handle stronger

loads. The propulsion system will also have 4 additional thrusters, which will add

complexity to the support structure. We size the frame the same way we size the frames

in the other designs. We then choose to scale the frame by a factor of 4 to account for the

added complexity. The complexity arises from extra structural support needed to mount

and harness the additional thrusters into the frame securely. We finally add another 15

percent of structural mass to account for our findings in Section 5.3.1. We found that the

frame was not strong enough based on our analyses and needed to add approximately 15

percent of additional structural mass to satisfy requirements. See Section 5.3.1 for details

on the configuration and layout of the support assembly and Section A-7.3.1 for complete

details on cross section dimensions.


Author: Tim Rebold

Electronics Module

The electronics module remains relatively similar to the other two missions with the

addition of scaling up structural members to handle stronger loads. See Section 5.3.1 for

details on the configuration and layout of the support assembly.

Skin

The skin will remain at 0.5 mm and cover the same areas of the OTV as in the previous

two missions. The increased size and volume of the OTV will drive up the surface area

of the skin. See Section 5.3.1 for details on the configuration and layout of the skin

Integration

Unlike the previous two designs, the large payload OTV has the same diameter as the

Lunar Lander. As a result, the LL skirt is not as complex. We use a scaling factor to

determine the size and mass of both Payload Attach Fitting (PAF) and LL skirt

integration components. See Section 5.3.1 for details on the configuration and layout of

the support assembly and Section A-7.3.1 for the new LL skirt and PAF masses.

Sizing

We use the same basic principles that we used to size the OTV in the 100g payload

mission for the large payload mission. See Section A-5.3.1 for complete details on this

process.



Material Selection

The large payload case orbital transfer vehicle is made of many of the same materials that

the 100 gram and 10kg OTV‟s are fabricated from. The floor and shear panels are the

only difference, being fabricated from aluminum 6065 in this case for extra strength and

support. All other structural members are made of aluminum 6065 as well. All bolts,

rivets, fasteners, and connectors are made of AIS 1015 low carbon steel.

FEA Analysis

Unlike the previous missions, we cannot perform a finite element analysis (FEA) for the

large payload mission, because we do not have enough time to model a new Orbital

Transfer Vehicle (OTV). The OTV for this mission is significantly different from the

previous mission configurations. Therefore, there will be no FEA analysis for our large

payload design.




Trajectory

Using the analysis explained in Section A-5.3.2 we designed the path for the large

payload. Table 7.3.2-1 summarizes the parameters of the trajectory.

Table 7.3.2-1 Large Payload Trajectory Configuration

Parameter Value

Payload Mass (kg) 3542

Thrust (mN) 2118

Mass Flow Rate (mg/s) 93.5

Earth Phase Angle (deg) 216

Moon Phase Angle (deg) 270

Parking Orbit Altitude (km) 400

Capture Orbit Altitude (km) 25

Initial Mass (kg) 9953

Flight Time (days) 365

Fig. 7.3.2-1 depicts the trajectory for the large payload.

Fig. 7.3.2-1 Large Payload Trajectory.

(Kara Akgulian)


Author: Levi Brown

Lunar Capture Orbit

The lunar capture orbit altitude is 25 km. This is the same altitude used as in the 10g

payload case. Please refer to Section 5.3.2 for more information.


One requirement for this project includes the capability to perform a 50 m/s burn for

course correction. Due to factors beyond the scope of this project, such as Sun

perturbations, the spacecraft will deviate from our trajectory design during actual flight.

To account for this bias, we carry additional propellant for the main engine. The

propellant is available for making small corrections in flight as necessary. We calculate

the propellant necessary for a 50 m/s burn and record the results in Table 7.3.2-2.


Parameter Value

Isp (s) 2250

mo (kg) 9953

Propellant for Correction (kg) 22.5


Author: Brad Appel


Propulsion System Overview

We select a cluster of five High-Power Hall Effect Thrusters (HPHETs) as the primary

propulsion for the OTV. Each thruster is substantially larger than the one used in the

100g and 10kg missions, however its supporting propulsion system is fundamentally the

same. Figure 7.3.3-1 shows the general layout of the EP system on the Orbital Transfer

Vehicle.

Fig. 7.3.3-1 Propulsion System general layout inside the Orbital Transfer Vehicle.

(Brad Appel)

We design the EP system to deliver cargo from Low Earth Orbit to Lower Lunar Orbit at

the cheapest price per kilogram. The main system parameters are listed in Table 7.3.3-1.

Table 7.3.3-1 Propulsion System Totals


Wet Mass 3,810 kg

Dry Mass 520 kg


Burn time 365 days

Payload Capability 4,545 kg

PPU

FCS

Thruster


Author: Brad Appel

One composite overwrapped tank provides the Xenon storage for all thrusters. The Flow

Control System is slightly more sophisticated than in the other missions, since we branch

out the feed lines into a five different paths. Each branch contains essentially the same

components found in the 100g and 10 kg missions.

Each thruster is a Busek BHT-8000. We optimize the combination of specific impulse,

mass flow rate, and number of thrusters based on cost. Table 7.3.3-2 provides the

operating conditions of each of the five thrusters.

Table 7.3.3-2 Specifications for the BHT-8000 Hall Thruster


Thrust 424 mN



Power Input 7,600 W

Efficiency 0.64 --

Mass 25 kg

By using five thrusters we also increase the system reliability. We include enough margin

in the required power so that even if one of the thrusters fails, the remaining four could

complete the mission. Also, we have the ability to apply a controlled torque to spacecraft

by throttling particular engines. Here, throttling means adjusting the power input, not the

mass flow rate.

We achieve thermal control in a similar fashion to the 100g and 10kg missions. Even

though the waste heat out of the thrusters totals to almost 14 kW, there is enough surface

area in the bottom skin of the spacecraft for passive radiation control.


Author: Brian Erson

7.3.4 – Attitude

Attitude Control

As mentioned in Section 5.3.4, three systems are integrated to perform attitude control of

the OTV.

1. Sensors – The same set of sensors are used for all payloads: a sun sensor and star

sensor, both made by VFCT

2. Reaction Wheels – Due to the increased mass of the OTV, a larger reaction wheel

must be used to maintain attitude stability throughout the translunar phase. VFCT

manufactures the VF MR 19.6, a wheel capable of creating adequate torque for the

system.

3. Thrusters – The 𝐻2𝑂2 thruster hardware needed is the same as the 100g and 10kg

case. An increase in inert mass does occur as shown in Table 3 below


complete system accomplishing the mission of controlling the OTV during the trans lunar

phase. Below, Table 7.3.4-1 provides an overview of our ACS specifications for the large

payload.

Table 7.3.4-1 OTV ACS Budget for Large Payload

Device Mass (kg) Cost ($) Power Required(W) Volume (m^3)

VF STC 1 (star sensor) 6.4 133,333 20.4 0.0163

VF SNS (sun sensor) 0.7 133,333 5 0.0163

VF MR 19.6 42 133,333 280 0.0258

H202 thruster 0.36 2,500 -- 0.00003

H202 Propellant 10.6 100 -- --

Inert Mass 13.84 1,000 -- 0.00957

Totals 73.9 404,000 305.4 0.068


Author: Brian Erson


The large payload case presents a significant increase in propellant mass. Although the

thruster hardware is the same, much more propellant must be used to create enough

torque for attitude stabilization. The increase in propellant yields the need for larger and

more massive tanks. This increase is not large enough to merit a change in propellant.

𝐻2𝑂2 still fits the bill as a powerful and adequate propellant for the large payload.


Reaction wheels are used to counteract the attitude perturbations due to the space

environment. Refer to Section 5.3.4 for further details.





For the large payload we will be using the same systems as for the 100g payload. Refer to

Section 5.3.5 and Section 5.4.5 for more information.

Communication Link Budget

Please see Section 5.3.5 for the communication link budget during the lunar transfer. The

communication requirements and equipment are unchanged with the change in vehicle

and payload mass.


For the arbitrary payload the ground station usage is the same as the 100g payload and

can be found in Section 5.3.5.

Communication Antenna Pivot

The patch antenna pivot discussed in Section 5.3.5 will be used again for the large

payload case, and will be essentially unchanged.





Similar to the 100g and 10kg payload cases, for the large case we thermally control the

electronics, the xenon tank, and the electric thruster. The electronic components put out a

total of 4189W of heat. To cool these components, we make use of the same thermal

control system as in the 100g and 10kg cases. We see the mass and volume breakdown

of the thermal control components for the electronics in Table 7.3.6-1.


For the large payload case, we also employ a simple wire heater to keep the xenon within

its storage temperatures.


The electric thruster generates approximately 16523W of heat that we need to dissipate to

keep the thruster under its maximum operating temperature of 473K (200ºC). We see the

dimensions of the electric thruster thermal control system as well as the rest of the OTV‟s

thermal control system in Table 7.3.6-1.

Table 7.3.6-1 Large Payload OTV Thermal Control Dimensions



Heat Pipes 15.43

ID = 3.04cm; OD = 3.34cm;

Length = 5m

Radiators 23.4

Total Cross-Sectional Area =

17.31m2

(8 fins @ 2.163m2 each)





7.3.7 – Power

Introduction

For the large payload case, we cannot use the Ultraflex-175 solar arrays that we

employed with the 100g and 10kg cases. We instead use rectangular solar arrays that can

deploy to a larger area.

Power Budget

We see the adjusted power budget for the OTV during the translunar phase in Table

7.3.7-1. The largest contributing factor to the OTV budget still arises from the electric

thruster. This device consumes over 90% of the entire OTV power budget. To ensure

that we meet all of the power needs, we increase the total power production by 5%.

Table 7.3.7-1 Large Payload OTV Power Budget by Group

Group Power Units




Power 1731.5 Watts

Lunar Lander

(during translunar) 105 Watts

Total 42960 Watts

Solar Array Sizing

With the power budget, we determine the sizes of the solar arrays. For the large payload,

we use rectangular arrays. The radii of the circular arrays are limited by the length of the

vehicle, and as such, we cannot physically attach large enough circular arrays to the side

of the OTV to support the power needs. Rectangular arrays offer more area for the same

length since they can deploy out several times their length. This OTV needs

approximately 43kW kilowatts of power, which can be provided by two arrays each with

dimensions of 5m x 14.3m. The solar arrays are composed of triple-junction gallium-



arsenide photocells. We see the dimension of the mass and deployed area for two arrays

in Table 7.3.7-2.

Table 7.3.7-2 Large Payload OTV Solar Array Dimensions (total)


Mass 286.4 kg


Similar to the 100g and 10kg payload cases, one of the largest driving factors in

determining the OTV‟s mission configuration for the large case is the high cost of the

solar arrays. The solar arrays cost a total of $42.96 million.

Battery Sizing

For the large payload case, we make use of the same lithium-ion battery cells that we use

for the 100g and 10kg cases. We see the new dimensions of the battery in Table 7.3.7-3.

Table 7.3.7-3 Large Payload OTV Battery Dimensions



Volume 0.0957 m3


Although the OTV‟s battery is still made of the same types of lithium-ion cells, we have

222 cells instead of ten or 13 cells. This results from the increased battery energy

capacity for the large case.

The components for the OTV‟s power system for the 100g and 10kg payload cases are

also used for the large case. We reference Fig. 5.3.7-2 in section 5.3.7 for a diagram of

the OTV power system components.



Table 7.3.7-4 Large Payload OTV Power Subsystem Dimensions



Battery 271.7 0.0957

PCDU 194.8 0.3437


Mission Configuration – Large Payload – Lunar Descent Section 7.4.1,

Author: Ryan Nelson



We determine the mass of the frame structure for the large payload case by scaling up the

100g payload case structure. The Lunar Lander for the large payload case scales from the

100g case because the design of the Lunar Lander does not involve two extra hopper

engines as with the 10kg payload case. These extra hopper engines force the height of

the Lunar Lander frame to be larger. The scaled size of the propellant system (main

engine nozzle, chamber and propellant tank) compares to the 100g payload case. Table

7.4.1-1 compares of the frame dimensions and scale factors between the 100g payload

frame and large payload frame. The frame mass for the large payload case is 104.86 kg.

Table 7.4.1-1 Scale up of Large Frame Dimensions from 100g case

Structural Component

100g

Dimension

Large

Dimension

Scale

Factor

Diameter of Floor Outer Ring 1.3 m 3.6 m 2.77

Diameter of Engine support Ring 0.315 m 0.872 m 2.77

Length of Rectangular Floor supports 0.493 m 1.364 m 2.77

Length of Side supports 1.011m 2.022 m 2

Diameter of Top Ring 1.0 m 2.4 m 2.4

Length of Legs 0.607 m 1.214 m 2.4

The side supports of the Lunar Lander for the large payload case scale up by a factor of 2

while the floor components of the Lunar Lander frame scale up by a factor of 2.77 as

depicted in Table 7.4.1-1. The top ring and legs increase by a factor of 2.4. The 3.6 meter

diameter Falcon 9 payload fairing constrains the Lunar Lander base. We use a scaling

factor of 2.77 to fit the base within the faring. Scaling the other aspects of the Lunar

Lander on such a large scale is unnecessary because this leads to empty space within the

Lunar Lander. As the amount of unused volume in the Lunar Lander increases, the mass

needed for thermal control also increases. With these scaling factors, the total volume of

the Lunar Lander is 14.33 m3.


Author: Ryan Nelson

It is impractical to design the basic Lunar Lander frame for the large payload case the

same as the Lunar Lander frames for the 100g or 10kg payload cases. With the previous

two cases, the small Lunar Lander frame sizes result in light frame masses without

getting too complex. Designing the frame for the large case in the same way the frames

are designed for the 10kg and 100g payload cases results in extremely thick frame

components and unnecessarily large frame masses. This is primarily because the four

rectangular beams running from the floor‟s outer floor ring to the engine floor support

ring experience high bending loads. Figure 5.4.1-1 represents a schematic of this design.

In order to get a more accurate depiction of the Lunar Lander frame mass for the large

case Lander_Frame_hover.m is run using the dimensions of the 100g Lunar Lander.

This yields a mass for each of the Lunar Lander components that are then scaled by the

respective factors seen in Table 7.4.1-1.

Material Selection

The large payload case Lander is made of many of the same materials that the 100g and

10kg Landers are fabricated from. The floor and shear panels are fabricated from a

magnesium alloy, while all other structural members are made of aluminum 6065. All

bolts, rivets, fasteners, and connectors are made of AIS 1015 low carbon steel.

CAD

The CAD model for the large payload Lander is designed the same manner as the Lander

for the 10kg payload. The only differences are that the large payload Lander is scaled up

to accommodate the larger payload and the large payload Lander does not need the two

small hopper engines. If you are interested in the design and layout of the internal

systems please reference Section 6.4.1.




Overview

In Section 5.4.2, we described the descent profile of the Lander for the 100g payload.

The descent profile of the large payload is very similar. Differences arise in the thrust of

the descent main engine, burn times, amount of propellant used, and perilune altitude

where final descent begins. In this case, descent begins at an altitude of 12.5 km. Also

note that the Lunar Lander is brought to a stop at 100 m where the hover trajectory takes

over. Tables 7.4.2-1 and 7.4.2-2 highlight the additional differences for the 10 kg

payload.

Table 7.4.2-1 Final Descent Burn Times


Radial Burn 000 - 169 34500

Rotation 169 - 174 3450

Vertical 174 - 188 34500



Mass (kg)

De-Circularize

9.5

Radial Burn 000 - 169 1952.0

Rotation 169 - 174 5.7

Vertical 174 - 188 161.0

Unusable

66.4

Contingency

12.0

Total

2206.6

Footnotes: An additional 79.6 kg of propellant

is needed for the hover, bringing the total

propellant mass in LPO to 2286.2 kg.

Landing Location

The landing location is in Mare Cognitum, which is the same location as the 100g

payload. To read more about the landing site please refer to Section 5.4.2.




Overview

The Lunar Lander main engine we designed for the large payload case is nearly identical

to the main engines used on the two smaller payload cases described in Section 5.4.3.

The only configuration change we make, is increasing the number of fuel grain plates,

shown in Fig. 7.4.3–1. The additional fuel grain plates allow us to optimize the

combustion chamber dimensions and corresponding mass.

Fig. 7.4.3–1 Radial flow hybrid engine dimensions and configuration.

(Thaddaeus Halsmer)


This engine is a scaled up version of the engine discussed in Section 5.4.3 and the only

performance parameters that change are thrust and burn time.


Author: Thaddaeus Halsmer/Saad Tanvir

Table 7.4.3 -1 Hybrid Engine Performance Specifications


Thrust Max / Min 34500 / 3450 [N]

Burn Time 188.9 [s]


The propellant feed system architecture is unchanged from the system described in

Section 5.4.3.


Because the propellant feed system is scaled up for this payload case, the power

requirement for the fluid control components is 275 watts during operation.

H2O2 Storage

Employing the same technique as in Section 5.4.3, we determine the amount of H2O2

decomposed over the period of one year. The hydrogen peroxide is stored in a sealed

spherical Aluminum 1060 tank. The amount of H2O2 decomposed during the mission is

2.29 kg. As compared to the 2,290 kg of total propellant carried, this value is negligible.



the H2O2 tank.




Volume 1.652 m3

Thickness 3.376 mm

Diameter 1.467 m

Mass 63.87 kg


Author: Saad Tanvir


the mission (1.588 m3: including the usable propellant, boil-off volume and the volume of

unusable propellant) and a 4% ullage volume (0.0635 m3).



the material for the He tanks (Humble). Table 7.4.3-4 shows the detailed specifications

of the pressurant tank as well as the mass of the Helium gas required of the whole landing

mission.




Volume 0.756 m3

Thickness 1.78 cm

Diameter 1.13 m

Mass of the Tank 25.49 kg

Mass of the Helium 26.08 kg


The inert mass of the propulsion system obtained as described in Section 5.4.3. The total

propulsion system inert mass for the hybrid landing engine equals 277.18 kg. Table 7.4.3-

5 gives a detailed breakdown of the propulsion system inert masses.


Author: Saad Tanvir






Nozzle 115.50 kg


Injector 2.60 kg


Feed System 2.60 kg


Valves 6.94 kg


Thermal Analysis

As in Section 5.4.3, similar thermal analysis is performed for this case. We found that to

keep the H2O2 at 283 K temperature during the lunar transfer phase we require a constant

power supply of 37 W.

During lunar descent, the temperature drop in the hydrogen peroxide tank is 2.3 K (from

283 K to 280.7 K). Similar temperature drop from the valves and the feed lines are 2.2 K

and 3.7 K respectively. Because the drop in temperature of these components during

lunar descent is small, its affect on engine performance is negligible. Hence, there is no

power required for thermal control on the propulsion system during lunar descent.




Much larger attitude control devices are required for this payload case due to the increase

in descent vehicle size. Due to the fact that the attitude sensing devices are not mass

dependent the same Valley Forge attitude sensing devices will be used for the large

payload case. In order to combat the larger thrust that the main lunar descent engine

creates the attitude control thrusters must be larger as well.

The attitude thrusters are aligned in the same fashion as the previous lunar descent

modules and are composed of twelve 335 N General Kinetics hydrogen peroxide attitude

control thrusters. As in the previous two cases we use the hydrogen peroxide from the

main lunar descent engine tank for attitude control (General Kinetics). The use of the

attitude control thrusters are for a brief de-orbit burn and then for the descent phase to a

height of 100 meters above the lunar surface.

The de-orbit burn using the attitude control thrusters will require 9.5 kg of hydrogen

peroxide. For the descent burn, which occurs until the vehicle is 100 meters above the

lunar surface, we use the attitude control thrusters to control the rotation of the space craft

due to torque created by main engine offset. The thrusters are oriented in the same

fashion as in the previous payload cases and are fired in opposing pairs to offset

detrimental torques and to insure there will be no translational motion from the attitude

thrusters. During the descent phase it will cost 92.9 kg of hydrogen peroxide attitude

propellant to counteract the torque we create with the lunar descent main engine.

Other mass concerns for the attitude system include the piping that we use to tap off the

main engine propellant tank and a laser altimeter used for determining the height of the

spacecraft above the lunar surface. With all components factored in the total mass of the

attitude control system is 125.38 kg for the descent portion of the large payload mission.





For the arbitrary payload the communication equipment onboard the Lander is the same

as the 10 kg payload and can be found in Section 6.4.5.


Please refer to Section 5.4.6 for details regarding the Communications Power

Requirements for the Lunar Lander/Hopper.


Please see Section 5.4.5 for the communication timeline and GLXP requirements for the

large payload case, as they have gone unchanged and the way that we will achieve them

is unchanged as well.




Lunar Day

We use the other two payload cases‟ thermal control system to design the large payload

case. The difference between this case and the other two is the locomotion. The case does

not need the extra thermal control due to the Space Balls or the hopper engines. This

design includes only one main engine and oxidizer tank needing thermal control. Figure

7.4.6-1 is the diagram of the system.

Fig. 7.4.6-1 Schematic of Passive Thermal

Cooling System.

(Josh Elmshaeuser)

The differences in this payload‟s thermal control system are the heat pipe‟s length and the

components needing a multilayer insulation (MLI) blanket or heater. The heat pipe‟s

length increase is due to the increase in the Lunar Lander‟s area. The heat pipe extends

from five feet in the other cases to twelve feet in this payload case. The MLI blanket

covers the Lunar Lander, main engine, and oxidizer tank. The main engine, oxidizer tank,

and the radiators require a heater. The total power during Lunar Transfer needed for



thermal control is 21 Watts. The other systems scale with the increase in the surface area.

The breakdown of the masses is found in Table 7.4.6-1. The total mass of the thermal

control system is 56.54 kilograms. The total mass was calculated in the MATLAB

program LanderThermalControl.m.

Table 7.4.6-1 Day Thermal Control of Lunar Lander


MLI Blanket 21.4

Aluminum Plate 1.40

Heat Pipe 10.53

Radiators 22

Ammonia 0.1727

Heaters 1.03

Total 56.54

Lunar Night

We will encounter the same environmental conditions for the large payload configuration

as with the 100 g payload. The most pronounced difference is the greater surface area of

the protected enclosure so that the heat loss is approximately 119 W instead of 11 W.

We are using the same heating and shutdown system as described in Section 5.4.6. The

main difference is the amount of hydrazine is increased by a factor of 11 as is the tank

volume to compensate for the increased heat loss.



7.4.7 – Power

Battery

We size the battery for the Lander based on power requirements from the Lander engine,

communication and attitude systems. The Lander engine, the communication gear and the

attitude systems all require power from the battery during the landing phase.

The Lander‟s engine requires power to operate the various valves and pumps to make the

engine function. A total of 275 W are required and the engine operates for 450 seconds.

The communication gear uses a peak total power of 54.53 Watts, to ensure that we

maintain constant communication between the Earth and the Lander throughout the entire

Landing phase. We provide enough power for the communication gear to transmit for 30


Watts to power our system and requires this peak power for 15 minutes to make sure the

Lander arrives at its destination. Table 7.4.7-1 highlights the total power requirement for

all systems and their individual components on the 100 g Lander.

Note: Complete description of each device can be found in

Section 5.4.2

For a complete background on each device within each system, please refer to tables A-

6.4.7-1 through tables A-6.4.7-3 in Section A-6.4.7.

We purchase the batteries from the company Yardney Technical Products, Inc., which

provides commercial off-the-shelf lithium ion batteries that are used in space vehicles.

We select a 21 Ampere-hour battery to support the operations of the Lander.


Item Power Units



Attitude 25.4 Watts

Total 354.93 Watts



Solar Array Sizing


week long lunar day. Refer to Section A-6.4.7 to see breakdown of power requirements

for each device within the communication equipment.

Because the power requirements for the Lander do not change for the large payload case,

the design and size for the solar cells is the same as for the 100 g and 10 kg cases. Figure

7.4.7-1 displays the power consumed and the power generated by the cells through a

lunar day.

Fig. 7.4.7-1 Maximum Potential Power of solar cells during a lunar Day.

The communication gear has enough power to meet its peak power

requirements on the 1.5 day of the lunar day. At maximum power, the solar

cells provide 74.5% more power than required at day 7 of the lunar day.

(Adham Fakhry)

From Fig. 7.4.7-1, the cells still provide enough power for the Lander during its entire

time on the lunar surface. Since the solar cells remain the same, the cost of the solar

arrays is the same and is listed in Section 5.4.4.



Solar array sizing is calculated in lander.m. Refer to Section A-5.4.7 for breakdown of

costs and the constants used to size the solar cells.

DC-DC Converters

Table 7.4.7-2 displays the DC-DC converter system for the large payload.



Mass 0.985 kg

Dimensions 0.033 x 0.033 x 0.033 m


Cost $ 68,500 -

Note: Cost is in 2009 US dollars and is the cost for all the units.





Please refer to Section 5.4.7 for the complete analysis of why we need a PCDU, what is

used for and its specifications.

Mission Configuration – Large Payload – Locomotion Section 7.5.1,


7.5 – Locomotion

7.5.1 – Hover Trajectory

The requirements for the arbitrary payload locomotion phase are to travel 500m on or

over the surface of the Moon using a propulsion system on the Lander to perform a

“hop”. However, for this payload the Lander does not have to touch down on the Moon‟s

surface before beginning its locomotion. We choose to travel the 500m by using a

combination of the main lunar descent engine and the attitude control system to perform a

translational hover and landing. The main lunar descent engine will be throttled to control

the Lander‟s vertical movement while the attitude control system will control the

Lander‟s horizontal movement.

The Lander comes to a complete stop at an altitude of 100m above the lunar surface

before the hover trajectory starts. Then the attitude control system fires to move the

Lander horizontally while the main engine throttles down to begin the descent. During

the last part of the trajectory, the attitude control system fires in the opposite direction to

stop the horizontal movement while the main engine throttles up to stop the Lander‟s

vertical movement. This trajectory saves propellant and time compared one where the

Lander touches down first.

Figure 7.5.1-1 shows the hover trajectory and Table 7.5.1-1 provides the results of this

trajectory.



Fig. 7.5.1-1 Lander hover trajectory.

(Alex Whiteman)

Table 7.5.1-1 Hover Trajectory Results


Time of Flight 61.4 sec

Propellant Used 19.9 kg

Average Thrust 1011 N

-700 -600 -500 -400 -300 -200 -100 0 100-50

0

50

100

150

range (m)

altitu

de

(m

)

Moon

Hover Trajectory




Overview

The Lunar Lander main engine continues to operate during the Locomotion Phase using

its variable thrust capability to perform the hover, locomotion and landing maneuvers.

The Lunar Lander main engine used for locomotion is fully described in Section 7.4.3.


For the large payload case the Lunar Lander main engine is operating continuously

throughout the Lunar Descent and Locomotion phases. Average thrust and burn time

during the Locomotion phase is given in Table 7.5.2–1.

Table 7.5.2-1 Engine Performance Specifications


Thrust (average) 3954 [N]

Burn Time 61.4 [s]


The propellant feed system architecture is unchanged from the system described in

Section 7.4.3.


During this phase of flight the power requirement is unchanged from what was given in

Section 7.4.3.




For translating the lunar descent module the required 500 meters we will use the attitude

control thrusters. The thrusters are oriented in a fashion that allows us to maintain a hover

condition with the lunar descent main engine. It can be seen in Fig. 7.5.3-1 that the

thrusters oriented in the cluster of four and control the rotation about the b3 vector are

able to induce translational motion when fired in pairs with the thrusters on the opposite

side of the vehicle.

At the end of the 500 meter translation the control thrusters are used to rotate the craft

and orient the main engine thruster in a direction that will allow for a soft landing. The

benefit of using the attitude control thrusters for the translational motion is that it will

allow us to adhere to the requirement that the hover phase must take over one minute.

The amount of hydrogen peroxide propellant the attitude thrusters use from the time the

craft reaches its hover altitude at 100 meters until touchdown on the lunar surface is 3.6

kg.

Fig. 7.5.3-1 The above figure shows the orientation

of the thrusters about the lunar descent vehicle. It

should be noted that the thrusters that control the

rotation about the b3 axis are able to cause

translational motion when fired in the proper

pairs. Not to scale.

(Josh Lukasak)

b3

b2

Mission Configuration – Large Payload – Vehicle Integration Section 7.6,



Much of the integration process for the large payload case would be similar to the 100

gram and 10kg payload cases. In this case, though, the integration was only broadly

touched on. We simply did not have the time to go into great depth on where things were

going to be placed in this vehicle due to the fact we only have one semester to complete

the project. With two other missions being concurrently integrated, that left little time to

accomplish great detail in integrating this mission. For the most part, components present

on the 100g and 10kg OTV‟s simply scale up to accomplish the large payload mission.

With a similarly scaled up frame, plenty of room exists within the body to place the

components and still have an OTV that fits in the Falcon 9 fairing.

One notable difference in the mission integration of the large payload case is the lander

and OTV now have the same diameter, and a mating skirt is no longer required.

Once this was completed, we integrate the entire set of vehicles in the same manner as

the 100 gram case and the mission is ready to launch.

Mission Configuration – Large Payload – Risk Analysis Section 7.7,




success for each vehicle in the mission. We compute the probability of success for the

large payload case using a slightly modified version of the method used for the GLXP-

sized payloads. We assume that higher-cost missions tend to have a higher probability of

success than low-budget missions. We increase the probability of success for each

system under the assumption that a larger integration expense tends to increase the

reliability of the system.

Unlike the GLXP-sized payload cases, the large payload does not include a separate

locomotion phase. Reusing the lunar lander for locomotion increases the probability of

success for the mission. A table describing the probability of success for each vehicle is

shown below.



Falcon 9 Launch Vehicle 99%


Lunar Lander/Hover 97%


This mission success rate satisfies the 90% mission requirement in one launch. Our large

payload mission can reliably deliver payload to a lunar base.

Mission Configuration – Large Payload –Cost Analysis Section 7.8,




Expense Cost ($M)

Falcon 9 Launch Vehicle 38.2


Lunar Lander / Hover Vehicle 28.6

Overhead 49.1


$/kg Payload $128k

Our costing method tends to underestimate the total mission costs. This total cost is

difficult to compare with historical values. Our team could not locate total mission cost

estimates for the Surveyor or Ranger programs. We believe the Surveyor program is an

ideal comparison to this mission, as their mission requirements are similar and have

comparable masses.

Due to lack of information, we are forced to make a cost comparison to the Apollo

program. The Apollo program, a manned lunar mission, is at least an order of magnitude

more complex than our unmanned mission. In addition, the mass of the Apollo descent

stage is roughly three times the mass of our descent stage, reducing the overall accuracy

of the comparison. Using the marginal cost of an Apollo mission, the cost per kilogram

of payload to the lunar surface is approximately $8M per kilogram, adjusting for

inflation. This is many times the cost of our mission. However, we do not believe this is

a fair comparison to our mission.

Alternative Designs Section 8.1.1,


8 – Alternative Designs

Alternative Designs Section 8,


Introduction

Throughout the design process, a number of alternative designs were considered. In

order to maintain transparency, we feel it is important to include these alternatives in the

main body of the report. The following section provides an overview of some of the

important trades considered in our mission configuration. As the majority of the

alternatives we considered were relatively independent of payload size, this alternative

section is organized by mission phase.

Alternative Designs – Launch Vehicle Section 8.1.1,


8.1 – Launch Vehicle Alternative Designs

8.1.1 – Launch Vehicle Alternatives

Prior to our final selection of launch vehicles we gave consideration to many other

vehicles. We evaluated each vehicle based on the launch cost to LEO for the smaller

payload cases and cost per kilogram to LLO for the large payload case.

Table 8.1.1-1 Small Launch Vehicle Characteristics (Futron Corporation)

Vehicle Payload to LEO-200 km (kg) Cost per kg to LEO ($/kg)

Taurus (USA)

Start (Russia)

Falcon 1e (USA)

Kosmos 3M(Russia)

Rockot (Russia)

Long March 2C (China)

Strela (Russia)

Long March 2D (China)

Dnepr (Russia)

1380

632

1010

1500

1850

2400

1700

3500

4400

13,768

11,687

9,100

8,667

7,297

7,031

6,300

5,000

4,800

Table 8.1.1-2 Medium Launch Vehicle Characteristics (Futron Corporation)

Vehicle Payload to LEO-200 km

(kg)

Cost per kg to LEO ($/kg)

Atlas 2 (USA)

Ariane 44L (ESA)

Delta 3 (USA)

Long March 2F (China)

Long March 2E (China)

Soyuz (Russia)

Falcon 9 (USA)

8640

10200

8291

8400

9200

7000

10454

11,029

11,314

10,000

6,700

5,500

5,400

3,600

Alternative Designs – Launch Vehicle Section 8.1.1,


Table 8.1.1-3 Large Launch Vehicle Characteristics

Vehicle Payload to LLO (kg) Cost per kg to LLO ($/kg)

Delta IV M (USA)

Delta IV M (5,2) (USA)

Long March 2C (China)

PSLV (India)

Delta IV M (5,4) (USA)

Delta IV M(4,2) (USA)

Long March 2E (China)

Delta IV H (USA)

Atlas V (USA)

Dnepr (Russia)

Ares V (USA)

Falcon 9 (USA)

Falcon 9 H (USA)

6360

7705

1517

2206

9874

8900

4466

18545

7836

2108

80564

6780

22234

26,536

24,680

23,568

22,470

20,595

19,798

17,288

17,251

16,733

12,521

6,788

6,782

4,973

Selection

We select the Dnepr-1 for the small payload missions because it yields the lowest launch

cost. We select the Falcon 9 for the large payload mission because it yields the least

expensive cost per kilogram to LLO. Although the Falcon 9 Heavy configuration

outperforms the medium lift version, the amount of thrusters needed is not feasible for

our Orbital Transfer Vehicle.

Alternative Designs – Lunar Transfer Section 8.2.1,

Author: Levi Brown

8.2 – Lunar Transfer Alternative Designs

8.2.1 – Trajectory Alternatives

Bielliptic Transfer

While performing preliminary analysis on various methods for lunar transfer, we consider

a bielliptic transfer. A bielliptic transfer involves three impulse burns for a total transfer

of 360o. Figure 8.2.1-1 created in Satellite Toolkit (STK) shows an example of the

transfer shape.

Fig. 8.2.1-1 Example of Bielliptic Transfer.

(Levi Brown)

Burn 1 accelerates the spacecraft from the initial parking orbit (periapsis of 1st transfer

arc) to some intermediate radius r (apoapsis of 1st transfer arc). Burn 2 accelerates the

spacecraft again, so the resultant periapsis of the 2nd

transfer arc equals the final orbit

radius. For a lunar transfer, the final orbit radius is the Moon‟s semi-major axis of

384400 km. Burn 3 decelerates the spacecraft to circularize the orbit at the final orbit

radius.


Author: Levi Brown

We perform an analysis of a bielliptic transfer and compare the results to a Hohmann

transfer. As we see in Table 8.2.1-1, the bielliptic transfer requires more ΔV, which

consequently results in a larger initial OTV mass. We conclude that for a transfer

employing chemical propulsion, a Hohmann transfer is the most cost effective method.

Table 8.2.1-1 Bielliptic vs. Hohmann Transfer Result Comparison

Earth Parking

Orbit

km

Lunar Parking

Orbit

km

Intermediate

radius

km

Parameter Hohmann Bielliptic

200 110 1 x 106 ΔV (km/s) 4.0 4.2

TOF (days) 5 81



Weak-Stability Boundary

A weak-stability boundary transfer saves 25% ∆v when compared to a Hohmann

Transfer. The savings can nearly double the amount of allowable payload placed into

low lunar orbit (Belbruno). This transfer begins with the OTV exiting the Earth capture

orbit and performing a fly by maneuver with the Moon. Enough energy has been gained

from the Moon‟s gravity so that the vehicle can now reach the weak-stability boundary of

the Earth. At this location the OTV performs another maneuver to travel back to the

Moon and be ballistically captured. We must align and time all bodies correctly to

successfully execute this trajectory. This analysis would be a massive project that would

span beyond the scope of this feasibility study. Because of the complexity of the weak-

stability boundary transfer and its highly theoretical background it does not make this a

viable option to perform.



Hohmann Transfer with Lunar Capture

The first trajectory that we consider for the tranlunar phase is the well-known Hohmann

transfer. A Hohmann transfer consists of two impulsive engine burns, typically carried

out by chemical propulsion systems. The Hohmann transfer and lunar capture is

illustrated in Fig. 8.2.1-1. We note the second ΔV is a braking maneuver to capture into a

circular lunar orbit.

Fig. 8.2.1-1 Earth-Moon Hohmann transfer generated using Satellite Tool Kit (STK).

(Andrew Damon)

ΔVpc for a simple plane change maneuver is also calculated. All ΔV values are tabulated

in Table 8.2.1-1.

ΔV1 ΔV2



Table 8.2.1-1 Gravitational and Orbital Parameters for the

Earth-Moon Hohmann Transfer


ΔV1

ΔV2

ΔVtot

ΔVpc

3.131

0.820

3.952

2.166

km/s

km/s

km/s

km/s


Author: Levi Brown

Trajectory Alternative Selection

We compare the methods described in this section to determine the optimum mission

architecture.

We find that weak stability boundary poses several problems in the design process. The

method is primarily theoretical, so easily making changes for different payloads becomes

troublesome. Based on the design cost to employ this method, we discard it.

Comparing a Hohmann and bielliptic transfer, we see that the Hohmann consistently

requires less ΔV and time. No matter the configuration, initial OTV mass is less for a

Hohmann transfer.

We compare the required initial OTV mass for a Hohmann and spiral transfer. We find

that the ΔV for a Hohmann transfer is approximately 4.0 km/s. Previously, we calculated

the initial OTV mass that results for this ΔV for different propellants. Additionally we

find the initial OTV mass taking advantage of a spiral trajectory (See table 8.2.3-1).

Table 8.2.3-1 OTV Mass Comparison

Transfer Type Initial OTV Mass (kg)

Hohmann 992

Spiral 550

We see that the initial OTV mass is significantly less for a spiral transfer, which

consequently leads to a lower cost. We select a low thrust spiral transfer as the mission

architecture.


Author: Saad Tanvir

8.2.2 – Propulsion Alternatives

Chemical Propulsion Alternative

In the case of chemical propellants, we size the Orbital Transfer Vehicle (OTV) using the

ideal rocket equation and historical values of finert, Isp and oxidizer to fuel ratio. A ΔV of

4 km/s is assumed for the Lunar Transfer phase. Table 8.2.2-1 shows the OTV propellant

masses and the total OTV wet masses using different chemical propellants.

Table 8.2.2-1 OTV wet masses for different chemical propellants

Propellant OTVwet mass (kg) OTVwet Propellant mass (kg)

LH2/LOx 992 594

LCH4/LOx 2084 1470

RP-1/LOx 2350 1681

MMH/N2O4 2514 1812

Solid 4690 3550

Based on the results shown in Table 8.2.2-1, we see that LH2/LOx is the most feasible

chemical propellant option. However, due to high complexity, cost as well as mass

compared to electrical propulsion system, the chemical option was retired.



EP Gimbal

In an effort to decrease the mass of the OTV, we considered many propulsion

alternatives. A notable alternative among those is adding a main thruster gimbal. The

basic concept is to give the OTV the capability to move and point its main thruster so that

some attitude control and maneuvering can be done without engaging attitude control

thrusters. The primary benefit of this alternative is the potential to decrease the attitude

control propellant which, in turn, decreases the total mass and the total cost of the

mission. The main disadvantage of adding a gimbal mount for the main engine is that the

mount introduces extra complexity, another failure point, and adds mass. According to

Vaughan (2005), a low complexity 3-axis gimbal mount could be developed, built, and

tested in house for less than about $1000 and with about 6 kg of material. This system is

shown in Fig. 8.2.2-1 with the OTV geometry and has a maximum range of about 20

degrees.

Fig. 8.2.2-1 This is a representation of the OTV

equipped with a 3-axis gimbal mount on the main

thruster. It also includes the location of the

center of mass and critical OTV dimensions at the time

the alternative was considered.

(Kristopher Ezra)



At the stage of development when we considered this alternative, it was noted that

onboard reaction wheels (spinning wheels which orient the OTV by pointing their

angular momentum vectors in a specific way) could completely manage attitude control.

Because of this, the only remaining task is de-saturating the reaction wheels periodically

(about 6 times) throughout each day so the wheels can continue to function properly.

Since the attitude control thrusters are located on the external body of the OTV, they

necessarily have a longer moment arm and produce more torque with less thrust. It was

this distance and increased torque that showed, after some analysis, that it was more

costly in terms of mass to gimbal the main engine than to use attitude control thrusters for

de-saturation instead. In all, the de-saturation propellant mass required by the attitude

control thrusters was found to be approximately 6.3 kg whereas the de-saturation

propellant mass required by the main engine was found to be 17.9 kg (We credit Brian

Erson with the development of the desaturation propellant cost model). This large mass

difference drives our decision to fix the main engine and use attitude control thrusters to

de-saturate reaction wheels.


Author: Brad Appel

Lunar Transfer Propulsion Selection

Early in the design process, we see that the electric propulsion system is capable of

delivering the same payload to the moon while using less than half the mass of the

chemical propulsion system. The cost of this mass advantage is a long time of flight and a

large power requirement, but we can still accomplish the mission in less than a year. The

EP system offers further advantages because it requires fewer moving parts, takes up less

space, and is easier to obtain than the LOx/LH2 chemical propulsion system. Therefore,

we find that electric propulsion is the best choice for a cheap and simple cargo mission to

the Moon.



8.2.3 – Attitude Alternatives

Hydrogen Peroxide vs. Hydrazine Attitude Control Thrusters

We need the OTV to have the ability to de-saturate (or de-spin) the reaction wheels

during the Lunar Transfer phase. The most common method of performing these

maneuvers is with small attitude control thrusters. These thrusters could be useful for

attitude control during lunar descent as well (about 20Nm is needed for the 100g and

10kg payloads). One of the proposed main propulsion methods is to use hydrogen

peroxide (H2O2) as an oxidizer, so it may be possible to utilize this for attitude control as

well. Valves make both hydrazine and H2O2 throttleable (Marotta, Circle Seal Controls).

As we show in tables A-8.2.3-1 and A-8.2.3-2, hydrogen peroxide thrusters are both less

massive and significantly more affordable than hydrazine thrusters (Astrium, Rocket

Research Corporation, Wernimont, Astronautix).

Electrothermal Hydrazine

A summary of the characteristics of two types of electrothermal hydrazine thrusters,

resistojets and arcjets, is shown in Table 8.2.3-1. We note that the alternative hydrazine

monopropellant system uses less than one watt per thruster. Based on cost, power

consumption, and complexity of system integration, neither of these options was selected

for our mission (Delft, Sellers, Pinero).

Table 8.2.3-1 Resistojet and Arcjet Thruster Data

Criterion Resistojet Arcjet Units

VacuumThrust 0.5-1 0.1-0.23 N

Input Power per thruster 500 300 W

Maximum Isp 300 500 sec

Life Span 380 1250 hours

System Mass (4 thrusters

+ PPU) 6.5 1.22 kg

Note: PPU is Power Processing Unit; Cost is about $150,000 for either option


Author: Tony Cofer

8.2.3 – Power Alternatives

Introduction

We consider several different options for providing power to the OTV during the Lunar

Transfer phase. Based on the power system designs of past successful spacecraft, we

analyze four alternatives: nuclear radioisotope thermoelectric generator (RTG), fuel cells,

batteries, and solar arrays.

Nuclear Power

Nuclear power on spacecraft has been used primarily on missions past the orbit of Mars

where limited solar power is available, or where there are extended periods of darkness.

Power outputs for these systems have always been less than 1 kW electrical.

Radioisotope thermal generators or RTG‟s transform thermal energy from natural

radioactive decay into electricity by means of thermocouples that are only about 10%

efficient (Angelo, 1985). We would then have nine times as much waste heat to remove.

With our 2500 watt system this would mean 22,500 watts of heat energy that would have

to be removed which would require increased mass for large radiators negating any

benefit.

Fuel Cell

For long duration missions fuel cells are often used as the primary power source. They

have advantages of long working life and a high energy density with a steady output.

Disadvantages are their expense, complexity, and need for volatile cryogenic fluids.

Another possible disadvantage of using fuel cells on this mission could be maintenance.

Fuel cells have to be purged of contaminants twice daily and the products of the chemical

reactions need to be expelled as they would be dead weight (Dumoulin, 2009).

Space-rated fuel cells have chemical efficiencies of only about 10% so it is impractical to

use them for our electric propulsion system when efficiencies of greater than 25% are

obtained by using the fuel in a chemical rocket engine (Patel, 2005).



Primary Batteries

Primary batteries are batteries that cannot be recharged because of non-reversibility in the

chemical reaction undergone to produce electrical power. The only portion of the mission

that these types of batteries are useful for are portions of the mission that do not include

solar panels. The only phase that did not use solar panels is the locomotion phase which

contains the Space Ball and the most mature model of the Rover. All other portions of the

mission contain the need to recharge batteries and therefore primary batteries are not

considered in these phases. The criteria for selecting the Space Ball primary battery is

described in section A-5.5.5.

Secondary Batteries

Secondary batteries are batteries that have a chemical composition that can be reversed

and can be recharged after it is depleted. This makes them necessary for any portion of

the mission that uses solar panels as the primary source of power. There are several

chemical make-ups that can be recharged and are shown in table A-8.2.4-1.



Solar Array Power

One of the major advantages that solar arrays have over other forms of energy production

is their minimal storage volume. We can fold and stow solar arrays during times of

inactivity (launch) and can deploy the arrays when they are needed (translunar phase).

Current solar array technology is also very resistant to temperature changes and radiation

damage. These two characteristics allow us to stow the solar arrays outside of the

spacecraft, providing volume inside of the spacecraft for other critical systems. This also

reduces the mass of insulation we need to protect the power system from the harsh

environment of space.

Solar arrays are ideal for spacecraft whose power demands are less than 10,000W and

whose mission durations are greater than a few days (Griffin, 2004). For the 100g and

10kg payload cases, our power requirements for the OTV are less than 3kW; this makes

our design a good fit for solar arrays. In addition, because of the long duration of our

mission, we also select solar arrays for the large payload case. Solar arrays are also very

reliable due to their successful integrations with numerous past and present spacecraft.

Table 8.2.3-1 Solar Cell Technology Comparisons

Cell Type Efficiency W/m2 $/Watt (2008) Maturity

Silicon 12-14% 150 112 In Use

GaAs/Ge 18-19% 225 337 In Use

GaInP2/GaAs/Ge 30% 300 337 New & In Use

Amorphous Silicon 8% 80 100 In Development

CuInSe 10% 135 100 In Development

Note: This table was developed using data from Patel

There are several different types of solar cells from which we can fabricate our solar

arrays. These options are detailed in Table 8.2.3-1. Because of their wide availability

and relatively low costs, silicon-based solar cells have been most commonly used in past

spacecraft, including the International Space Station. These types of cells are limited in

capability, however, due to the fact that they are low in efficiency, relatively

susceptibility to radiation-induced degradation, and have a higher mass per power output



than other types of cells. The efficiency of silicon solar cells is around 13%, which we

see from Table 8.2.3-1 is less than half as efficient as other types of cells (Patel, 2005).

From this fact, we see that in order to equal to the power output of other types of solar

cells, the area of the silicon arrays would need to be twice as large as the area as other

more efficient cells. It is evident to us that these characteristics would increase the mass

and the subsequent launch costs associated with silicon cells. Based on the high launch

cost per kilogram for this mission, we determine that the savings in purchase costs would

be well offset by the additional launch costs from increased mass. From the drawbacks

of silicon-based solar cells, we easily determine that silicon cells, while common on past

spacecraft, are not suitable for this mission.

Based on the properties in Table 8.2.3-1, we select triple-junction gallium-arsenide solar

cells for the OTV‟s solar arrays. These cells are very efficient because they are

constructed with three layers of sunlight-absorbing materials. These materials consist of

GaInP2 (gallium-indium-phosphorus), GaAs (gallium-arsenide), and Ge (germanium).

Not only are these solar cells significantly more efficient than silicon cells, but they are

much thinner than silicon cells (~0.25mm, including protective coverglass). Gallium-

arsenide cells are also much more resistant to adverse changes in efficiency at higher and

lower temperatures. The cells are also much less susceptible to radiation-induced

degradation (Patel). All of these characteristics greatly contribute to reducing the overall

required mass for the solar arrays.

From Table 8.2.3-1, we see that the largest disadvantage to the triple-junction solar cells

is the high cost per watt of power. These cells are approximately three times as

expensive as the silicon-based cells. These costs, however, are offset by the two-fold

increase in power generating capabilities. This characteristic allows the solar arrays to be

smaller in size and, subsequently, smaller in mass than the silicon arrays. The savings in

launch costs due to this reduction in mass significantly offset the increase in price per

watt of the solar cell.

Alternative Designs – Lunar Descent Section 8.3.1,


8.3 – Lunar Descent Alternative Designs

8.3.1 – Landing Alternatives

High Energy Tangent Landing

The High Energy Tangent Landing (HETL) alternative involved the Lander impacting

the Lunar surface with a significant portion of its orbital velocity. We wanted to explore

this design because it would lower propellant requirements for a soft landing. The

MATLAB file tangent_landing.m computes the horizontal distance our Lander will

slide with a given horizontal velocity and mass. In order to compute this distance some

basic assumptions were used:

1) The Lander would not skip.

2) The Lander would not dig into the Lunar soil.

3) The Lander would not be completely destroyed upon impact.



Figure 8.3.1-1 shows the horizontal slide distances of our Lander that were computed.

Fig. 8.3.1-1 Horizontal slide distance of Lander for three landing g-force cases.

(Trenten Muller)

Besides the large slide distance one other concern came about during this analysis. With a

significant portion of its orbital velocity our Lander would more likely create a crater and

be destroyed upon impact. Because our mission involved certain goals after landing this

risk would not be acceptable.

This design was not chosen because of this risk and because of the large slide distances.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

5

10

15

20

25

30

horizontal velocity (km/sec)

dis

tance (

km

)

10g

15g

20g



Gas Spring Landing Gear

Our vehicle orbiting the moon has high kinetic energy. The Lunar Lander engine

dissipates this energy during descent. Some of this energy could be stored during landing

and used for lunar surface locomotion as illustrated in Fig. 8.3.1-2. We examine gas

springs consisting of a piston and cylinder. The springs compress at touch-down and

later release the stored pressure to bounce the Lunar Lander along the lunar surface.

Preliminary analysis shows our 500 m locomotion could be achieved with bouncing.

However, gas springs have temperature and acceleration limits that are exceeded during

lunar landing (Enidine, 2009). Gas springs are unsuitable for landing gear because of

these violations.

Fig. 8.3.1-2 Schematic of Lunar Lander hopping with gas spring landing gear.

(Christine Troy)



Accordion

The communication equipment was found to take the least amount of force on the Lunar Lander,

10 g‟s. We find the forces acting on a three-dimensional object may cause deformation, causing

additional stress forces within the object. Newton‟s second law of momentum leads us to find

that the longer the contact time, the less the force on the object will be. Therefore, the greater the

deformation of a material, x, the less amount of force exerted on the object. Using a honeycomb

material at the bottom of the Lunar Lander would increase the amount of material will deform

while not adding as much mass if it was just solid. We can see an example of the honeycomb

core in Fig. 8.3.1-3.

Fig. 8.3.1-3 Honeycomb core schematic.

(Caitlyn McKay)



The assumption for the force exerted on the Lunar Lander is there would be complete

deformation of all hollow areas in the global buckling of the core. For the following

information the Lunar Lander used in the MATLAB code theAccordion.m, has a dry

mass of 230.8 kg excluding the material being deformed, radius of 1 meter, and a plate

face thickness of 0.001 meters. The results of the optimal honeycomb core type we see in

Table 8.3.1-1.

Table 8.3.1-1 Honeycomb core length and mass of material crushed at different velocities for a

force of 10 Earth G’s

Honeycomb core 1 m/s 5 m/s 10 m/s 15 m/s 20 m/s

3/8-5056-.0007p 0.0252 m 0.1283 m 0.512 m 1.151 m 2.050 m

1.25 kg 6.45 kg 25.74 kg 57.87 kg 102.84 kg

From the data we determined that the mass added to the Lunar Lander for the crushable

core to slow impact time was greater than that of the fuel that would be saved. Other

materials could have been looked into, but a concern would be that the material would

have to be strong enough to withstand the 8.5 Earth G‟s of launch. The Lunar Lander

could be supported during launch, but that would add more complexity and weight to the

orbital transfer vehicle. Even if the mass of the core was not a problem, the fact that the

crushable material has to be below the engine is a nuisance, not to mention that the

material would only buckle globally once which would not work for the 10 kg payload

vehicle. It was decided that incorporating shocks into the legs of the Lunar Lander would

be simpler than dealing with these problems.



Airbag Landing

We consider using inflatables or shock absorbers as an alternate landing method because

of the potential for mass savings over a soft thruster landing. We investigate an airbag

system comprised of six cylindrical airbags arranged circumferentially on the base of the

OTV, as shown in Fig. 8.3.1-4.

2.0 m

Fig. 8.3.1-4. Arrangement of six airbags on the base of the OTV.

(Andrew Damon)

We assume that the base of the OTV is 2.0 meters, and that each airbag is a cylinder of

length 0.5 meters. The airbags are evaluated under an impact of 10 times Earth gravity

acceleration (10 g‟s), which is equivalent to a deceleration of 98.1 m/s2. If we assume

the mass of the combined lander and locomotion phases is 250 kg, then the force on the

airbags is equivalent to 24525 N. If this force is divided evenly among the six airbags,

each airbag must withstand a force of 4087 N. The proposed airbag would be made of

Kevlar, with a valve or flap designed to yield at a specified pressure in order to prevent

the lander from bouncing. This “burst valve” is shown on the drawing of a single airbag

in Fig. 8.3.1-5.



Fig. 8.3.1-5. Side view of a single Kevlar airbag with burst valve.

(Andrew Damon)

Using a density value for Kevlar of 1440 kg/m3, we determine that an airbag thickness of

2 mm will withstand the shock of landing. The length of the airbag is 0.5 m, and the

diameter is 0.3 m.

We next determine the mass of the system in order to evaluate it against the additional

propellant mass needed for a soft landing. Six of the airbags described above will have a

total mass of 5.03 kg. In order to inflate the airbags, gas generators are necessary. The

estimated mass of gas generators needed to inflate airbags of this volume is 10 kg.

Miscellaneous structural mass necessary to contain the uninflated airbags and support the

gas generators is estimated at 5 kg.

Total Airbag System Mass: 20 kg.

Burst valve

0.3 m

0.5 m



Momentum Transfer

A preliminary alternative proposed for landing a payload on the lunar surface was the

idea of momentum transfer. In this scenario, the dry OTV traveling at orbital velocity

(approximately 1.7 km/s) ejects a portion of its mass toward the moon. By conservation

of momentum or a work-energy construction, we expect that the portion of the dry OTV

containing the Lander slows by some amount to allow it to make a semi-soft landing on

the lunar surface. In this analysis it is assumed that mass is ejected in one piece by a

spring-like force and any energy lost is attributed to this ejection. Based on preliminary

analysis, it happens that this approach is entirely infeasible.

To apply a mathematical model, we assume that a reasonable collision (or explosion)

distance is on the order of 1-2 meters. This short distance implies that, to dissipate the

full orbital velocity of 1.7 km/s, the Lander must be subjected to 1x105 Earth gravity

accelerations. This high acceleration is achieved when the dry OTV ejects the most

possible mass away from the Lander and so it is the minimum value we expect to be

reached over such a short collision distance. The limiting factor is the communication

equipment which can withstand an acceleration of at most 10 Earth g‟s. At this low

value, the collision/explosion required would have to act over a length of 150 km to slow

the Lander to a stop at the lunar surface. Figure 8.3.1-6 demonstrates the acceleration felt

by the Lander is subjected to versus the collision distance required. These accelerations

are tremendously high for short range collisions. Since the accelerations and forces are

probably greater than what is sustainable by the molecular bonds in the Lander, we

observe that a new approach to landing is probably necessary and discard the momentum

transfer alternative.



Fig. 8.3.1-1 Plot of the acceleration that the Lander is subjected to versus the collision distance

required to slow the Lander from 1.7 km/s to 0 km/s. These accelerations are in multiples of Earth's

gravity (g's).

(Kristopher Ezra)

0.5 1 1.5 2 2.5 3 3.5 4 4.5 50

0.5

1

1.5

2

2.5

3x 10

5 Work/Energy Based Accelerations

Collision Distance (m)

Accellera

tion (

g)



Semi-Soft Landing

We considered a semi-soft landing as an alternative to a soft landing to see if it would

save mass or cost on the lander. We determined a 15g maximum loading would

constitute a semi-soft landing. Initially, we thought the benefits of this landing would be

a savings in complexity and propellant mass. However, after looking at this alternative in

more detail, we determined that the propellant savings would be at most a few kilograms

unless the lander could impact the lunar surface with a significant percentage of its orbital

velocity. This scheme, however, would require a large addition to the structural mass of

the lander in order for it to withstand such high loading. Consequently, this alternative

was discarded in favor of a soft landing.



Spinning Tether

An alternative explored for placing our Lander on the lunar surface was the concept of a

spinning tether. In this alternative, the Lander and dry OTV are connected by a tether

and spin about their center of mass during descent. If the system spins so that the Lander

is traveling at the orbital velocity of 1.7 km/s around the center of mass, then when the

Lander moves toward the Moon it travels with a relative speed of 3.4 km/s and when it

moves away it has zero relative speed (this only occurs when the tether is exactly

perpendicular to the velocity vector of the entire system). To capitalize on this effect, we

ignite a charge exactly when the Lander is at or near the lunar surface and has zero

relative speed. The dry OTV is lost, but we place the Lander on the surface of the Moon

with little or no propellant. Figure 8.3.1-7 illustrates the setup. For simplicity, the figure

shows the center of mass of the system to be in the center of the tether, this is an extreme

simplification and is obviously not the case in the actual analysis.

Fig. 8.3.1-2 The system of the dry OTV with the Lander moves toward

the Moon at a rate v while rotating at a rate w. The goal is that the

Lander moves with a net speed of zero relative to the Moon at some

point along its trajectory.

(Kristopher Ezra)

We observe that this alternative is subject to the same constraints as the other landing

alternatives in that the maximum acceleration sustained by the Lander cannot exceed 10

v

w2v

v=0



Earth g‟s, the system must travel at 1.7 km/s initially, and for the system to be viable it

must have a mass less than the proposed propellant mass. With these constraints in mind,

we assume that the dry OTV and Lander are connected with a length of Kevlar (whose

safety factor is 1.25 in all cases observed) and use a work/force/energy model to recreate

the scenario. At the point in time this alternative was considered, the dry OTV had a

mass of 251 kg, the Lander had a mass of 138.3 kg, and the descent propellant mass was

75 kg. Fig. 8.3.1-8 shows the required length of tether for a 1.7 km/s orbital velocity is

about 50 km.

Fig. 8.3.1-3: Plot of the magnitude of the orbital velocity versus the tether length required to spin the

Lander at this rate without exceeding the 10 Earth g's constraint.

(Kristopher Ezra)

0 10 20 30 40 50 600

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

Tether Length (km)

Magnitude o

f Lin

ear

Velo

city (

km

/s)

Linear Velocity vs Tether Length



Given this tether length, we compute the net mass savings if the tether replaces the

descent propellant. The density of the tether remains constant and so does its cross

section (1.3 mm radius including a 1.25 safety factor), but as the length of the tether

increases so does its total mass. Fig. 8.3.1-9 shows the mass savings as a function of

tether length and it is clear to us that, even if we could support a tether with a length that

is half of the orbital height above the moon, the mass of the tether is too large. A mass

savings of -325 kg indicated on the plot corresponds to a tether with a mass 325 kg

greater than the descent propellant. This equates to a tether with a weight of 400 kg.

This is not feasible. For this reason, the spinning tether alternative was discarded.

Because of the very low accelerations sustainable by the communications equipment (10

g limit), the Lander and payload cannot be placed on the Lunar surface by means of a

spinning tether system.



Fig. 8.3.1-4: Total mass savings obtained when the tether replaces the descent propellant. The

function decreases with tether length and rapidly becomes negative.

(Kristopher Ezra)

0 10 20 30 40 50 60-400

-350

-300

-250

-200

-150

-100

-50

0

50

100

Tether Length (km)

Mass S

avin

gs (

kg)

Mass Savings



Alternative Selection

There are a variety of creative landing trajectories with which we could have chosen for

our mission but in the end the most elegant and safest landing trajectory is a soft landing.

Every landing trajectory explored aside from soft landing is an unproven method for

lunar landing. The lack of a lunar atmosphere makes any alternative except a soft landing

very risky and reduces our probability of mission success. The mass additions that are

required for the some of the alternative landing trajectories (spinning tether and

accordion) are simply just too costly to implement and explore. The structural mass

increases to protect the electronics on board the lunar lander during a semi-soft or hard

landing are far too large and complicated to use a landing trajectory. In summary the

reduction in mission success probably, implementation cost, and mass increases drove us

to choose the soft landing alternative for the lunar descent phase in all payload case

instances.


Author: Ryan Nelson

8.3.2 – Structural Alternatives

Structural Design

We use a conic frustum shape for the Lunar Lander in each of the payload cases for two

reasons. This shape allows for the comfortable housing of all the Lunar Lander

subsystems without leaving too much unused volume. This simple design shape also

allows for easy recalculation of the frame mass as the size and mass of the Lunar Lander

subsystems change.

We freeze the conic frustum shape itself in all payload cases, but alternative design

concepts for the frame around this shape are examined. The rectangular floor support

beams contribute the largest portion of mass to the Lunar Lander frame. The floor

consists of an outer circular ring connected to an inner engine support ring by these four

rectangular beams as depicted in Fig. 5.4.1-1. We examined the frame mass in all three

payload cases for three and five of these rectangular floor beams. We find that with three

of these rectangular floor beams each beam becomes thicker in order to support the

bending loads. The increased thickness causes an unnecessary increase in Lunar Lander

frame mass. With the examination of five rectangular floor beams, we find that although

each beam becomes thinner, the addition of the extra beam results in a higher Lunar

Lander frame mass.

A similar analysis was done on the side support beams which did not play a large roll in

the overall frame mass, but did contribute. Four side support beams are used for all three

payload cases but three and five supports are analyzed. The results compare with the

floor supports. For smaller payload masses such as the 100g and 10 kg cases, a Lunar

Lander frame with four side supports and four legs yields the lowest overall frame mass.

These alternate frame designs are not analyzed for the frame with the arbitrary payload.

Since the basic Lunar Lander frame design depicted in Figure 5.4.1-1 is impractical for

the frame with the arbitrary payload case, we used a scale up from the four side support

and four leg 100g payload Lunar Lander.



8.3.3 – Trajectory Alternatives

Linear Tangent Steering Law

Although our final mission descent trajectory employs a relatively simple radial/vertical

burn scheme, other descent trajectories were considered. The most promising of these

alternatives analyzed was the linear tangent steering law (LTSL), as presented to us by

James M. Longuski, Professor of Astrodynamics, Purdue University. The goal of this

section is to describe the analysis that was performed using the LTSL method, as well as

illustrate the cost-benefit conclusions of the analysis. Finally, we will explain why the

radial/vertical burn scheme was chosen for all three mission configurations.

The linear tangent steering law reduces analysis of a launch or descent to a two

dimensional problem. The variable of interest determined using the LTSL is the angle at

which a spacecraft is thrusting, here we use 𝜃, as a function of time. Figure 8.3.3-1

provides an overview of the problem; y is in the vertical direction pointing up from the

surface of the moon.

Fig. 8.3.3-1 Overview of descent from parking

orbit to landing, which is at the origin in this

figure.

(John Aitchison)

From Professor Longuski‟s presentation on the linear tangent steering low, for the Lunar

Lander we calculate the x and y components of position and velocity as a function of 𝜃,

𝜃0, a, and f. Please see the MATLAB codes LinearTangentSteeringLaw.m as well as

LinearTangentSteeringLaw_func.m for the process used to accomplish this.

x

g

y

𝜃

f



After using the LTSL to determine an optimized function for theta, the amount of

propellant used was compared to that required of our radial/vertical burn scheme. In each

case, the flight times using the LTSL were much greater than those obtained using the

radial/vertical burn scheme. These longer burn times using the LTSL resulted in higher

propellant masses based on a constant mass flow assumption. The LTSL did, in cases

where we started descent from very high altitudes, result in propellant savings. These

savings were seen when descent started altitudes above approximately 40 km.

As a result of the lower propellant mass usage of the radial/vertical burn scheme, we

chose not to use the LTSL. Although the LTSL did provide propellant savings when

starting descent at high altitudes, the total propellant usage from descent at low altitudes

trumped these savings. Also, a significant note is that the LTSL analysis assumed

constant acceleration of the spacecraft. This simplification would have significantly

reduced the validity of the descent trajectory analysis. Using the radial/vertical burn

scheme, we were able to model acceleration as a function of time using precise equations

of motion, which provided a more accurate model of the Lunar Lander descent


Author: Saad Tanvir

8.3.4 – Propulsion Alternatives

Bi-Prop vs. Hybrid Engine of Lander Engine

The use of a bi-propellant (LH2/LOx) system leads to a total system mass savings of ~6

kg for the 100 g and the 10 kg payload cases. However, no existing bi-propellant engine

exists that can satisfy the thrust levels that we require. Developing a new system from

scratch is infeasible for the mission time frame and extremely costly as compared to the

radial flow hybrid engine.

Furthermore, the bi-propellant system is more complex than the currently employed

hybrid system. If we use the bi-propellant engine, we will be adding another propellant

tank, increase driving gas volume for the pressurant system, increase the number of feed

lines and double the control valves. Moreover, power required for thermal control

approximately doubles and the overall volume occupied by the propulsion system

increases, which leads to an increase in Lander mass and volume.

As a result we eliminate the bi-propellant alternative for the Lander engine for all three

payload cases.



Solid Rocket Motors

Solid rocket motors (SRM) were one of the propulsion design alternatives we considered

for Lunar Descent. The advantage of this approach is the elimination of the relatively

complex propellant feed system associated with liquid propellants. By using multiple

solid rocket motors fired in series, analysis of the Lunar Descent shows that a semi–soft

landing is theoretically feasible. Depending on the impulse of the engines we find that

the complete Lunar Descent mission requires a total of 5–8 engines.

ATK was the only solid rocket motor supplier found that sold the desired size of motors.

When priced we found that the purchase cost of each engine was approximately

$700,000. Table 8.3.4–1 contains the motor types, cost and the number required to

perform the lunar descent mission using SRM‟s. The initial conditions for the analysis

were based on an initial descent altitude at the Perilune of a 110x15 km altitude elliptical

orbit. Because of the prohibitively high cost of this approach, further analysis of the

SRM lunar descent design was not pursued.

Table 8.3.4 – 1 SRM purchase costs for 85kg payload lunar descent

QTY ATK Solid Rocket Motor (STAR Series) Cost Each

4 STAR 13b (provide impulse to zero circular velocity) $700000

3 STAR 6b (provide impulse to zero vertical velocity) $700000

Total cost for lunar descent motors $4.9 million


Author: Brad Appel

Lunar Descent Propulsion Selection

The main incentive for using solid rocket motors for Lunar Descent is their simplicity –

but at a cost of several million dollars, we find it is worth it to go for a less-simple but

cheaper system. At the other extreme, a bi-prop descent engine requires an enormous

amount of complexity – and runs into the same problem as solids with commercial

availability. Monopropellant systems do not offer the competitive specific impulse

needed for a low-mass system. We find a happy medium with a radial Hybrid engine.

The Hybrid engine is relatively simple, offers good performance, and we can develop it

for a price cheaper than the other options.



8.3.5 – Attitude Alternatives

Spin Stabilized Lunar Lander

Much like a spinning top stands stably upright, a spacecraft spun around its central axis

remains fixed in space. Since space is a very low friction environment, once a vehicle is

spun up, it remains spinning will very little or no additional force input. For these

reasons, we consider spinning the Lunar Lander for attitude control. Spin stabilization

only requires propellant for initial spin-up and for spin-axis reorientation. Figure 8.3.5-1

is a schematic of the spinning Lunar Lander.

Fig. 8.3.5-5 Diagram of spinning Lunar Lander configuration.

(Christine Troy)

We calculate the propellant needed for both the spin up and axis reorientation and

compare this value to the propellant mass needed for 3-axis control. The mass of

propellant saved is only around 2.2 kg. There is also 0.54 kg mass savings because six

thrusters are eliminated. However, to implement the spinning Lunar Lander, the descent

engine and landing gear each need extensive redesign. These redesigns add significant

mass to the systems, outweighing the mass saved in propellant. Additional attitude

sensing equipment is also needed for the spinning Lunar Lander concept. The spinning

Lander concept is more complex and riskier than 3-axis stabilization and offers no mass

savings. For the above reasons, the spinning Lunar Lander alternative design is not used.



Attitude Control-Thrusters

The two main options for lunar descent attitude propellant are hydrazine and hydrogen

peroxide. An important factor when we choose attitude control propellant is the specific

impulse of the propellant. The specific impulse of hydrazine is approximately 230

seconds (Astronautix) while hydrogen peroxide has a specific impulse of approximately

140 seconds (General Kinetics, 2009). The reason that specific impulse is such an

important factor when determining what attitude propellant can be seen here:

(Rauschenbakh, 2003)

𝑚 = 𝑀

𝑔 ∗ 𝐼𝑠𝑝 ∗ 𝐿 (8.3.5 − 1)

In Eq. 8.3.5-1 𝑚 is the mass flow rate in kg/s of propellant. Our goal when using this

equation is to keep 𝑚 to a minimum, thus reducing mass and cost. When applying this

equation to the descent phase of the10kg payload case and comparing the results when

using hydrazine or hydrogen peroxide the mass savings for using hydrazine is 2.61 kg,

which is a fairly large amount. The issue with hydrazine though is the high cost of it as a

propellant and for thrusters that employ hydrazine. Hydrazine thruster systems cost

nearly $100,000 each while hydrogen peroxide thrusters cost on the order of thousands of

dollars. This cost savings alone is worth the extra 2.61 kg we have on the system due to

our choice of hydrogen peroxide. Not only the thrusters but hydrazine itself is in the

hundreds of dollars per kilogram for only the propellant while propellant grade hydrogen

peroxide is under ten dollars per kilogram (Peroxide Propulsion, 2009). The large cost

savings brought about by using hydrogen peroxide more than offsets the slight mass

advantage we gain by using hydrazine attitude control thrusters.

Alternative Selected

For the lunar descent module and all payload cases the attitude control system we chose

was hydrogen peroxide thrusters and 3-axis stabilized landing as opposed to spin

stabilized. In summary hydrogen peroxide thrusters cause a slight increase in mass on our

lunar descent vehicle but save us hundreds of thousands of dollars in system costs. 3-axis



stabilized landing is a less complex attitude control solution than spin stabilization. Spin

stabilization creates a reduction in attitude propellant requirements but the mass savings

would need to be translated into extra structures to deal with the spinning descent. For the

large masses we are dealing with spin stabilization is not an advantageous solution for

attitude control.



Alternative Selection

In previous sections various modes of attitude control were discussed and decisions were

made based on the feasibility of implication, cost, and any possible risks involved. The

first decision made for the orbital transfer vehicle is what type of attitude control thrusters

are to be used. When looking at various types of thrusters including cold gas, electric

propulsion, and hydrazine it was apparent that hydrogen peroxide thrusters are the

optimal solution for attitude control. In general hydrogen peroxide is a relatively cheap

propellant when compared to other attitude propellants. Cold gas thruster systems

themselves are quite expensive and electric propulsion attitude thrusters require a large

amount of power which we cannot divert from the main electric propulsion engine. We

can note that having hydrogen peroxide thrusters require their own separate tanks but the

price benefits of hydrogen peroxide for attitude control far outweigh the slight mass

increase in the overall system and the performance advantages of certain attitude

thrusters.

The second alternative for attitude control is spin stabilization as opposed to a 3-axis

stabilized spacecraft. The benefit of a spin stabilized space craft is a reduction in attitude

control propellant required. We chose 3-axis stabilization for our space craft during the

lunar injection phase because it is difficult, without adding considerable structural mass,

to spin the space craft but also keep the solar panels and communication antennas

properly aligned in their respective directions.



8.3.6 – Thermal Control Alternatives

Thermal- Active Control

We considered a mostly active thermal control system because past space missions

include active systems. Instead of a heat pipe, we could use a cryogenic cooler could

perform the necessary task. Upon doing research several different coolers would cool the

Lunar Lander. The problems associated with the coolers included high mass and power.

For example, a NGST HEC weighs 7 kilograms by itself, additional 5 kilograms for other

thermal control. The cooler requires 120 Watts to run (Donabedian, 2004).

Another form of active control, we considered using heaters to warm the Lunar Lander

during the lunar night. The idea is not ideal for our case due to the large amount of power

that heaters require. Our system saves over 2 kg, 120 Watts, and $30,000.



8.3.7 – Power Alternatives

Solar Array Deployment

Two different alternatives were considered for deploying the arrays after storage in the

Lunar Lander during descent. The active system includes a solar array deployment

system. The passive system includes a solar hinge. The passive system would add over 1

kilogram in deployment and the solar array deployment system would add 2.27 kilograms

(Davis, 2009). For both of these alternatives the solar arrays would need to be protected

during descent.

The current design for the solar arrays on the Lunar Lander attaches the arrays to the top

of the Lunar Lander. We determine this simple design to be the best alternative for our

particular mission. Both these options became invalid after the decision was made to

place the solar arrays on the top of the Lunar Lander. The advantages for our decision

include less weight and power. Also the system is more reliable since the solar arrays do

not need protection during descent, unlike storing in the Lunar Lander, and the ability to

collect power without the delay between landing and deploying.

Alternative Designs – Rover Section 8.4,


8.4 – Rover

Introduction

The Rover design, as the locomotion phase, comes from the Apollo‟s Lunar Rover, Mars‟

Rovers Opportunity and Spirit, and the new Rover prototype currently being designed by

NASA. Since NASA has consistently used rovers in the past, our first designs for the

locomotion stage are rovers devised for our purposes.

There were two different designs for the Rover. The first was to have a slower moving

Rover that would travel the 500 meters in 2 Earth days designed for the 100g and 10kg

payload cases. These Rovers had to be carefully thermally controlled as to not overheat.

The second Rover is designed to travel the 500 meters in 13 minutes and then overheat

for the 10kg payload case. The mass break down for the second designed rover, given by

the rover design team, we see in Table 8.4-1.

Table 8.4-1 Mass breakdown for the

second Rover design, 10kg payload case

Item Mass in kg

Battery 0.422

Antenna 0.1

Transmitter 0.21

CPU (computer) 0.2

Camera 1

Motor (4) 0.092

Mounts (8) 0.176

Gearhead (4) 0.03

Thermal Control 0

Magnesium Body 0.4

Space Blanket 0.58

Wheel (4) 0.58

Power (Extra) 1.92

Payload 10

Total 15.71

The Rover was not chosen for either for the 100 gram or 10 kg case because at that point

in time other systems of travel were found to be lighter, which made the entire vehicle

Alternative Designs – Rover Section 8.4,


system less expensive. When the Hopper, the current 10kg payload design, reached the

same maturity level of the Rover the cost and mass were very similar. We can conclude

that this means that there is more than one feasible solution to the problem.

Alternative Designs – Rover Section 8.4.1,

Author: Ryan Lehto

8.4.1 – CAD/Integration

We modeled the small rover for the 100 g and 10 kg payloads in CATIA. Figure 8.4.1-1

shows how the rover components are assembled. The four wheel drive assemblies are

mounted on the underside of the rover with aluminum motor mounts. We place the

electrical components including the camera, battery, CPU, and transceiver, inside the

rectangular rover body. The figure however does not reflect the true size of the battery for

the faster Kamikaze rover as design worked stopped before the CAD model was updated.

Fig. 8.4.1- 1 The system configuration for the small rover design.

(Ryan Lehto)

Wheel Drive Assemblies

Battery

Camera

CPU

Transceiver



8.4.2 – Communications

While the communication systems for the rover were not fully developed due to the

switching to the Space Ball design, the mass and size of many of the systems could have

been reduced.

We selected larger heavier radiation hardened equipment for the Rover. We also planned

on communicating with the Earth using S-Band frequencies instead of using the lighter

UFH systems and communicating to the Lander.

We originally selected a dish antenna for communication back to Earth. As well as a

transponder and a receiver which were powerful enough the send the signal all the way

home. In the end the Rover just like the communication gear it was equipped with were

upgraded and replaced with lighter, more efficient products.


Author: Ryan Lehto


System Overview

We analyzed a rover propulsion system that consisted of four individual drive assemblies

attached to the underside of the rover body seen in Fig. 8.4.1-1. Figure 8.4.3-1 shows

each drive assembly includes a D/C motor, a planetary gear head, and a wheel.

Fig. 8.4.3-1 Assembly of the drive system components

the D/C motor, planetary gear head and the wheel.

(Ryan Lehto)

Placement Consequences

The placement of rover drive assembly exposes the motors and gear heads to the

harshness of the vacuum of space. The motors we use are recommended for space

applications; however the drive assemblies must be lubricated with a special space rated

lubricant that will not leak out nor freeze in the extreme hot and cold of space.

Motor and Gearing

We chose a 19 mm diameter D/C motor with an accompanying 22 mm planetary gear

head. The motors share the load of the rover mass equally. The initial rover design

moved at a velocity of 0.01 m/s. We believed a larger velocity would be too risky as the

rover could become hung-up on an obstacle or overturned. Each D/C motor connects

Wheel Gear

Motor

0.076 m


Author: Ryan Lehto

with a planetary gear head to amplify the torque provided to the wheels. All the motors

combine to produce a torque of 0.85 N-m to power the rover. The rover turns similar to a

tracked vehicle by reversing two drive assemblies on the same side. The system allows

the rover to turn 360° in place for a turn radius of 0 m.

The wheels are designed to traverse the lunar terrain and absorb vibrations during roving

phase of mission. The space inside the wheel is filled with foam to prevent the

accumulation of lunar dust and debris. The average sized obstacle expected will be less

than 0.20 m, see Section 5.5.6. The optimum wheel size and spacing is 0.076 m 0.214 m

respectively. The curved spokes act as a dampener, absorbing the vibrations caused by

rover‟s movement. No further analysis of spoke effectiveness at absorbing vibrations

was performed due to the decision to use the Space Ball concept for the small payload.


Author: John Dixon

8.4.4 –Thermal Control

The proposed thermal control for the Rover Alternative was to use a solid copper block

heat sink. The copper heat sink would absorb heat from the electronics and transfer it to

heat pipes to conduct it to the surroundings. The primary focus would be to dissipate

heat from the processors of the electronic systems and the incoming solar energy. The

drawbacks of this system are the large weight of the copper heat sinks as well as the

inefficient dissipation.



8.4.5 – Power

The power system for the rover changed several times, most recent official number use a

single battery that outputs 3.6 volts and 12 ampere-hours. The total energy contained in

the battery cell is 43.2 Watt-hours. In addition to this 0.422 kg battery cell there is

addition hardware that pushed the total mass of the power system to 2.3 kg. The

additional hardware contained a kg of battery housing and DC-DC converters to increase

the voltage to the proper levels for each electrical component.



8.4.6 – Deployment

We use a linear shaped charge to deploy the Rover. When we ignite the charge there is

an explosive but straight cut through the material below. Linear shaped charges will be

made in-house using copper lining and C-4 as the explosive material (Kane, 2009).

Fig. 8.4.6-1 Linear shaped charge cross section.

(Caitlyn McKay)

We cut three sides of the Lunar Lander using the charge; a fourth side has a hinge that

opens. The Rover deployment system includes a platform in which the Rover is situated

that is lowered by a steel cable. An example of the deployment system is seen in Fig.

8.4.6-2. The lowering platform increases the complexity as well as the mass during

deployment as seen in Table 8.4.6-1.

Copper Lining

Explosive Charge



Fig. 8.4.6-2 An example of the deployment system used to lower the Rover from the Lunar

Lander to the surface of the Moon.

(Caitlyn McKay)

Table 8.4.6-1 Mass breakdown for Rover deployment

Item Mass in kg

Linear shaped charge 0.580

SOLIMIDE foam 0.030

Steel cable 0.82

Platform 0.025

Motor 0.025

Support Beams 0.13

Total 1.610 Footnote: The table is an estimate of masses.

Motor

Platform

m

Support Beam

Steel

Cable



8.4.7 – Structural Analysis

We initially designed the Rover using magnesium alloy as the material. Magnesium

while not as strong, has a lower density than aluminum. The magnesium alloy still

satisfied all of the Rover‟s stresses. For the second design of the kamikaze rover, we

were considering Lexan material. Since the Rover design was decided against, stress

analysis was not done on the Lexan Rover.

The stresses the Rover endures are caused by the forces exerted during different phases

and the mass of the systems inside the Rover. These forces cause bending moments,

deformations and buckling. Sheet buckling will occur if the support is not thick enough

and the Rover cannot withstand the compression, shear or bending stresses.

The design of the Rover wheel, seen in Fig. 8.4.7-1, is an airless wheel commonly

referred to as a twheel developed by Michelin. The wheel is made from aluminum with

chevrons added for traction and is filled with Solimide foam to keep dust and rocks out.

Fig. 8.4.7-1 Schematic of Rover wheel

(Ryan Lehto)

We designed the wheel size to increase the probability of a successful mission while

maintaining a low mass. Rocks the Rover would have to drive over were considered to

be steps that needed to be climbed. The optimum size for the base of the Rover is also

found by the radius of the wheel. The optimum base size allows the Rover to drive up a

45 degree angle.

Alternative Designs – Other Locomotion Section 8.5.1,

Author: Ryan Lehto

8.5 – Other Locomotion

8.5.1 – Sled Alternative

We investigated a tracked vehicle steered by skis as a means to traverse the lunar surface

for the small 100g and 10 kg payloads see Fig. 8.5.1-1. In our analysis, we assumed the

track vehicle to have same mass as the rover (see Section 8.4). We determined that 0.13

m by 0.017 m track is needed to power the vehicle.

Fig. 8.5.1- 1 Track vehicle schematic.

(Ryan Lehto)

The track system was not an attractive alternative as the advantages for the light payload

were minimal versus the disadvantages.

Advantages

Increases traction

Does not sink into soil

Disadvantages

Heavier than most other systems

Adds complexity

Kicks up lunar dust



8.5.2 – Spring Launch Alternative

For the purpose of locomoting our payload 500 meters across the lunar terrain we

investigated the feasibility of using a spring launch system. We calculated the potential

energy of different springs to determine the maximum distance the payload could

traverse. First, a few assumptions were made; launching from ground level, initially the

system is at rest, and a 45° initial trajectory angle to maximize range. The maximum

distance is calculated from basic dynamics.

We looked into launching our payload the full 500 meter distance to determine if this

design should be given further consideration. For the 100 gram payload case the spring

system could launch the payload over 1000 meters. However for the 10 kilogram payload

case a multiple spring configuration is needed to get the payload 500 meters away from

the base. For this reason the possibility of using a spring launch for the large payload case

is impractical.

We continued to look at the viability of a spring launch system for the 100 gram and 10

kilogram payload cases. The first calculation did not include the weight of a container for

the payload and other gear necessary to accomplish the mission; including the camera

and communication equipment. While developing the spring launch alternative the Space

Ball concept was introduced. The Space Ball has a shell that protects the payload and it

allows the equipment to travel across the lunar surface. However, the Space Ball could

not roll on a 45° inclined slope and therefore needed an external system to accomplish

this task. We decided that the spring launch concept would be developed concurrently

with the Space Ball design. At the time of calculation, the potential Space Ball design

weighed 7.6 and 18.5 kilograms for the 100 gram and 10 kilogram payload cases,

respectively.



Table 8.5.2-1 Spring Launch System Results

Variable 100 gram 10 kilogram Units

Mass

Free Length

Spring Constant

7.6

0.1524

18879

18.5

0.1524

18879

kg

m

N/m

Velocity

Number of Springs

41.6

30

40.7

70

m/s

--

Maximum Height 270 260 m

Distance 540 520 m

Table 8.5.2-1 shows that for both the smaller payload cases the spring launch design

completes the X PRIZE locomotion requirement. We determined the Space Ball shell can

withstand 30 g‟s upon impact; therefore necessary analysis was done on the g loads at

launch and landing. The results prove the spring launch system is incapable of delivering

the Space Ball a substantial distance away from the landing site because of the structural

limitations. The limiting factor of this concept is the g loads upon launch. We looked into

to the possibility of increasing the length of the springs to reduce that, however the

necessary spring length exceeds the scope of feasibility.

Table 8.5.2-2 Spring Launch System Results

Variable 100 gram 10 kilogram Units

Distance

Launch g‟s

Landing g‟s

Spring length for 30 g

launch

540

3549

26

18

518

3402

25.5

17

m

--

--

m

We considered launching a partial distance away from the lunar landing site rather than

the full 500 meter distance to reduce the length of the springs and the g loads upon

launch. The maximum distance the payload can launch, and still be within feasible limits,

is 200 meters. However, the requirement for the Space Ball to move up a 45° inclined

slope was relaxed therefore eliminating its need for an external system to achieve that

task. The added complexity of creating a system to store and un-store compressed

springs, which still do not accomplish the assigned task, is the final deciding factor for

the elimination of the spring launch system from project Xpedition.


Author: Cory Alban

8.5.3 – Ski Alternative

The ski has an advantage over a wheel based system in that, where wheels require

traction in order to provide forward motion, a ski actually benefits from low friction

environments. Instead of gripping the surface, it slides along the top providing a platform

of support. The large contact area of a ski distributes weight over a large area which

allows it to sit on top of fine materials (i.e. lunar regolith) without sinking in.

The ski is a natural spring which lets it act as a suspension system. The spring like quality

of the ski is helpful in dampening any vibrations created by moving across the lunar

surface.

The ski is not able to turn without carving an edge into the surface it glides on. A separate

motor system would be required to provide turning. The lunar regolith is twenty times the

density of snow thus requiring significant power required to turn. Though the ski would

slide over the regolith with low friction, a ski is still being pushed or dragged along. The

power required to overcome friction would need to be accounted for.

The ski system requires a propulsion source. Unless the Lander could land on top of a tall

hill, the ski system would still need an additional propulsion system. In the event that the

ski rides up onto a large obstacle, the chance of tipping the craft over or removing our

propulsion system from the ground is quite high. Due to these risks, we chose another

alternative for our locomotion design.

About The Authors Section 9,


9 – About the Authors

John Aitchison – As part of Mission Operations, John‟s primary responsibility was

design and analysis of the lunar descent trajectory. After graduation in May 2009, he will

be heading to Northrop Grumman Corporation in El Segundo, CA to begin work in the

space systems business division. Eventually he plans on starting his own company, either

in the aerospace or adventure travel industries, possibly both. His ultimate goal is to

work one day a week, and spend the other six kayaking the meanest whitewater in the

world and piloting commercial sub-orbital flights.

Kara Akgulian – Kara is a member of the Mission Operations technical group and the

Locomotion phase group. In May 2009 she will be graduating from Purdue University

with a major in design and a minor in dynamics and control. This summer Kara will be

getting married and plans to start working full time in the Aerospace Industry. In the

future she will continue her education with a Masters degree in Aeronautical and

Astronautical Engineering.

Cory Alban – Cory is a member of the Mission Operations group with most of his work

being done on the lunar locomotion phase. He will be working with Rolls-Royce in

Indianapolis this summer. Cory spends his time ballroom dancing, playing guitar, and

living the dream.

Brad Appel – Brad is the group lead for the Propulsion Group and focused on the

Electric Propulsion for the OTV. After graduation Brad will continue at Purdue with

graduate research in propulsion. He is currently scouting out living quarters in a van

down by the river.

Zarinah Blockton – Zarinah is a member of the mission operations group. She was

responsible for the launch vehicle selection for all three payload cases. She was also



involved in the locomotion phase of the mission. She researched the possibility of using a

spring launch system for the smaller payload cases. After graduation Zarinah plans to

teach overseas, over a year and a half period, in Vietnam and Chile. Upon her return to

the United States Zarinah will pursue a Master‟s Degree in Aerospace Design. Zarinah‟s

ultimate goal is to become an astronaut for the National Aeronautics and Space

Administration (NASA).

Levi Brown – After obtaining a Bachelor of Science in Aeronautical and Astronautical

Engineering, Levi will continue his education in the pursuit of a Master‟s of Science in

Aeronautical and Astronautical Engineering. He will primarily study propulsion and

astrodynamics. Upon receiving a Master‟s he will seek employment in the aerospace

industry. His long term goal is to become involved in trajectory analysis and mission

design for America‟s manned or unmanned space programs.

Michael Christopher – Michael Christopher is an undergraduate senior in the

department of Aeronautics and Astronautics at Purdue University. Originally Michael is

from Buffalo, NY. Michael‟s major contributions to Project Xpedition were in the

communications group. Upon graduation, in December of 2009, he plans on pursuing

employment most likely in the field of power generation. In the more immediate future,

during the summer of 2009 Michael will be interning for Solar Turbine in San Diego,

CA.

Tony Cofer – Tony hails most recently from Orange County Indiana and will graduate in

May. He is planning on going to graduate school at Purdue. His major is in propulsion

and minor in aerodynamics.

Andrew Damon – Andrew, hailing from Indianapolis, IN, will graduate in May 2009

with a BSAAE. He hopes to start a career working in manned spaceflight and also work



toward a graduate degree. Perhaps most importantly, he hopes to move out of the

midwest.

John Dixon – After finishing with my undergraduate education at Purdue University,

John will be attending the University of Nevada – Las Vegas to achieve a Masters in

Aerospace Engineering. Shortly after achieving the aforementioned degree he will then

begin his lifelong dream of World Domination. Falling short of that he will single

handedly invade Canada with nothing more than a bent paperclip, a piece of chewing

gum, a ball point pen, and a can of Pepsi, MacGruber style.

Joshua L. Elmshaeuser – Joshua‟s main contribution to project Xpedition was serving

as a member of the communications group. Within the communications group he

researched hardware options for the Lunar Transfer phase as well as the locomotion

phase. He served as the communications liaison for the translunar phase, and assisted the

locomotion group with some of the space ball design analysis. The best experience for

him was modeling the lander system in CATIA providing an encompassing perspective

of the spacecraft we are to put on the Moon. His future plans are to leave the field of

Aeronautical and Astronautical Engineering behind for a more hands on application in air

and space. Upon graduation he will be a Second Lieutenant in the United States Air

Force, stationed at Columbus AFB Mississippi training as a Pilot.

Brian Erson – Brian will be continuing at Purdue through the fall of 2009. After

graduation, he plans on pursuing a Masters of Business Administration, then help develop

marketing systems for engineering products.

Kris Ezra – Kris is a member of the Attitude Group working primarily in the Lunar

Transfer phase and focusing on landing alternatives. He has one more year of

undergraduate studies after which he will graduate with a bachelor of science in

astronautics with minors in applied physics, mathematics, and Spanish language and



literature. After graduation Kris will attend graduate school and will eventually pursue a

job with private defense contractors. He enjoys ballroom dancing, playing the piano, and

riding his motorcycle.

Adham Fakhry – Adham is a member of the Power group, focused on designing the

power systems for the Lander for all mission configurations and also worked on the

surviving the lunar night for the Lander. Adham is graduating this spring and hoping to

find an internship for the summer. After that he will be attending graduate school at

Purdue University to get his Masters degrees in Aeronautical and Astronautical

Engineering.

Thaddaeus Halsmer – Senior Design responsibilities were in the Propulsion group on

the Lunar Descent phase with primary focus on the Lunar Descent Vehicle main and

locomotion engine design. He will graduate in May 2009 with a Bachelor of Science in

Aeronautical and Astronautical Engineering and has accepted a position as a Rocket

Propulsion Development Engineer with Space Exploration Technologies (SpaceX) in

Hawthorne, CA.

Jeff Knowlton – Jeff plans to continue school until August when he graduates. His goal

is to become an astronaut, what path he will take to reach that goal is not set in stone, but

may include the military.

Kelly Leffel – Kelly worked on the Structures and Thermal Group for Senior Design.

Kelly spent most of the time working on the thermal control for the Lunar Lander. After

graduation in May, she will return to her hometown of Greensburg, Indiana to work a

summer job. Kelly plans on returning to Purdue University to receive a master‟s of

science in Aeronautics and Astronautics Engineering. After graduating with a master‟s,

she plans on seeking full time employment.



Ryan Lehto – After obtaining a Bachelor of Science in Aeronautical and Astronautical

Engineering with a minor in Computer Graphics Technology, Ryan will be working in

aerospace industry as a supplier technical support engineer contractor for Sikorsky

Aircraft. Long term goals include returning to school to obtain a Masters in engineering

and one day working as a test engineer for rocket propulsion firm.

Korey LeMond – Korey‟s main contribution to Project Xpedition was to serve as the

Structures Group Lead and as the systems engineer for integration of the vehicles and

missions. Korey will be graduating in May. Upon graduation he will be moving to

Houston where he will be working with Cameron International designing subsea oil

pipelines and wellheads.

Josh Lukasak – Josh was given way too much responsibility in this project and was

named Attitude Group Lead and Lunar Descent Phase Manager. Josh is graduating in

May ‟09 and is getting married in July because he‟s not as smart as you think.

Unfortunately for Josh the change promised by President Obama hasn‟t had a chance to

reach him before graduation. As a result he will follow his wife and live off her meager

teacher‟s salary while attempting to optimize burger patty trajectories or something just

as demeaning. All the while Josh will continue to brag that he went to the same

University as Neil Armstrong.

Caitlyn McKay – Caitlyn traveled from South Dakota to go to school at Purdue. She

was part of the Structures and Thermal group. This semester she focused on the

alternative rover design and was a 10 kg integration manager. She will commission as a

2nd

Lieutenant in the United States Air Force in May 2009 when she graduates. Caitlyn

will be stationed at Randolph AFB near San Antonio, TX to begin navigation training

starting one week after graduation.



Ian Meginnis – Ian served as the group leader for the Power Systems group and the

phase leader for the Lunar Transfer phase. He also worked on developing the thermal

control system for the orbital transfer vehicle. His major is design and his minor is

aerodynamics. After graduation, Ian plans on attending graduate school and then

working at NASA-JSC.

Trenten Muller – Trent was a part of the Communications group where he worked on

communication equipment as well as ground stations. He also explored an alternative

landing design that had the Lander impacting with a significant portion of its orbital

velocity. After graduating in May, 2009 he plans on joining the Air Force and working in

the aviation or space based field.

Ryan Nelson – This spring Ryan mostly did work on the frame design on the Lander for

each of the three payload cases. Early in the project he did some research and analysis on

the launch loads and vibrations for the Delta II launch vehicle as well as the Dnepr. After

senior design, he plan on graduating in May ‟09. Following graduation he plans on

attending graduate school in the Midwest and further expand on his understanding of

structures and propulsion as they apply to the aerospace industry.

Tim Rebold – Tim was part of the Structures team for project Xpedition, and worked

primarily on the OTV. He was responsible for designing the configuration of the OTV

and sizing its structural members. Tim is majoring in Dynamics and Control with a

minor in Structures while he attends Purdue University. He plans to continue his

education at Purdue University by seeking a Master‟s Degree in Aeronautical and

Astronautical Engineering. When not busy with school, Tim enjoys time with his wife

and daughter, and playing ultimate Frisbee and racquetball.

Saad Tanvir – Saad has primarily been working on developing the radial flow hybrid

engine for Lunar descent phase of the mission. After obtaining a Bachelor of Science in



Aeronautical and Astronautical Engineering, Saad will continue his education by

pursuing of a Master‟s of Science in Aeronautical and Astronautical Engineering. His

major area of research focuses on alternative energy with applications to the aerospace

industry. His long term plan is to return to Pakistan and help improve its aerospace

industry.

Christine Troy – Christine served as Project Xpedition‟s Assistant Project Manager,

Webmaster, and was in the attitude control technical area. Specifically, she worked on

attitude control for lunar descent. After graduation, she plans to work in either the

aerospace or wind energy industries.

Brittany Waletzko – Brittany is a member of the attitude control group and is also the

integration manager for the 100g payload for project Xpedition. After graduating this

spring, she is going straight into graduate school. Beginning with a math course over the

summer, she'll be working on getting her Master's degree in the area of Astrodynamics &

Space Applications at Purdue University's graduate school. Brittany expects the non-

thesis option will take less than two years to complete, after which she will be seeking

full-time employment.

Solomon Westerman – Solomon is the Project Manager for Project Xpedition. While

not devoting his time to Project Xpedition, Solomon enjoys cycling, guitar, and high-

power rocketry. He has accepted a Guidance, Navigation, and Control Engineer position

at Space Exploration Technologies (SpaceX) in Hawthorne, CA, where he plans to live it

up. You can check his website at http://web.ics.purdue.edu/~skwester and contact him at

[email protected].

Alex Whiteman – After finishing senior design, Alex will be graduating from Purdue

and then looking for a job. He also hopes to get a masters degree in Astronautical

engineering in the near future.

http://web.ics.purdue.edu/~skwester

References Section 10,


10 – References

1) About.com. “Composites and Plastics: Kevlar,” (2009). URL:

http://composite.about.com/od/aboutcompositesplastics/l/aa050597.htm [cited 1 April

2009].

2) Aerospace Technology, "EvonikFoams -SOLIMIDE Polyimide Acoustical and

Thermal Aircraft Insulation Foam," URL: http://www.aerospace-

technology.com/contractors/thermal/evonik/. [cited 1 March 2009].

3) Anaheim Automation, California, January 2009, "TGM42-075-40-12V-040A Series

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