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Project

Xpedition

Purdue University AAE 450 Spacecraft Design Spring 2009

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 6.6 – MISSION INTEGRATION ...... 184 6.7 – RISK ANALYSIS ...... 186 6.8 – COST ANALYSIS ...... 187 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 7.6 – MISSION INTEGRATION ...... 229 7.7 – RISK ANALYSIS ...... 230 7.8 – COST ANALYSIS ...... 231 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, Page 7

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

Author: Solomon Westerman Introduction - Background Section 2.1, Page 8

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.

Author: Solomon Westerman Introduction – What‟s in this Report? Section 2.2, Page 9

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.

Author: Solomon Westerman Introduction - Acknowledgements Section 2.3, Page 10

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

Author: Solomon Westerman Introduction - Acknowledgements Section 2.3, Page 11

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

Author: Solomon Westerman Introduction – Acronym List Section 2.4, Page 12

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

Author: Solomon Westerman Introduction – Acronym List Section 2.4, Page 13

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

Author: Solomon Westerman Project Overview – Design Requirements Section 3.1, Page 14

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.

Author: Solomon Westerman Project Overview – Design Requirements Section 3.1, Page 15

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.

Author: Solomon Westerman Project Overview – Design Requirements Section 3.1, Page 16

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.

Author: Solomon Westerman Project Overview – Design Requirements Section 3.1, Page 17

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.

Author: Solomon Westerman Project Overview – Interpretation of Design Requirements Section 3.2, Page 18

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.

Author: Solomon Westerman Project Overview – Design Process Section 3.3, Page 19

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.

Author: Solomon Westerman Project Overview – Design Process Section 3.3, Page 20

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

Author: Solomon Westerman Project Overview – Design Process Section 3.3, Page 21

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.

Author: Solomon Westerman Project Overview – Risk and Cost Analysis Section 3.4, Page 22

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.

Author: Solomon Westerman Project Overview – Risk and Cost Analysis Section 3.4, Page 23

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

Author: Solomon Westerman Project Overview – Risk and Cost Analysis Section 3.4, Page 24

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

Author: Solomon Westerman Project Overview – Results Section 4, Page 25

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.

Author: Solomon Westerman Project Overview – Results Section 4, Page 26

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.

Author: Solomon Westerman Mission Configuration – 100g Payload Section 5, Page 27

5 – Mission Configuration 100g Payload

Author: Solomon Westerman Mission Configuration – 100g Payload – System Overview Section 5.1, Page 28

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)

Author: Brittany Waletzko Mission Configuration – 100g Payload – System Overview Section 5.1, Page 29

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.

Author: Brittany Waletzko Mission Configuration – 100g Payload – System Overview Section 5.1, Page 30

Table 5.1-1 Mission Timeline

Elapsed Time Event Vehicle (ddd:hh:mm) -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 Take picture of Lander, Begin 000:00:15 Space Ball transmission to Lander 000:00:23 End photo transmission Space Ball Transmit arrival Mooncast (near real-time video, photos, HD video, 000:00:23 Lander XPF set asides, data uplink set) to Earth Transmit Mission Complete 001:33:56 Mooncast (near real time video, Lander photos, HD video) Finished transmitting, prepare for 002:08:04 Lander night 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.

Author: Brittany Waletzko Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.1, Page 31

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)

Author: Zarinah Blockton Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.1, Page 32

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

Variable Value Units Height 34.3 m Diameter 3 m Lift-off Mass 208,900 kg Mass to 400km orbit 3,400 kg Number of Stages 3 -- Orbit Inclination 50.5 degrees Cost per kilogram to LEO 4,800 $/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: Zarinah Blockton Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.2, Page 33

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: Andrew Damon Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.2, Page 34

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.

18

16

14

12

10

8

Launch and Lunar Transfer Cost(Million Transfer $) Lunar and Launch 6

4 400 500 600 700 800 900 1000 1100 1200 Parking Orbit Altitude (km)

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.

Author: Levi Brown Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.3, Page 35

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

Author: Kristopher Ezra Mission Configuration – 100g Payload – Launch Vehicle Section 5.2.3, Page 36

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.

Author: Kristopher Ezra Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 37

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 Axial Lateral Event Acceleration 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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 38

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:

퐴푙푙표푤푎푏푙푒 푦푖푒푙푑 푙표푎푑 (표푟 푠푡푟푒푠푠) 푀푆 = − 1 (5.3.1-1) 퐷푒푠푖푔푛 푦푖푒푙푑 푙표푎푑 (표푟 푠푡푟푒푠푠)

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 39

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 40

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 41

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: Tim Rebold Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 42

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 Yield Young‟s kg 3 ( /m ) Strength (Pa) Modulus (Pa) Al 6065 2700 1.03E8 6.8E10 Mg AZ31 1770 1.5E8 4.4E10 AIS 1015 Low 7800 2.55E8 2.05E11 Carbon Steel

Author: Korey Lemond Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 43

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 44

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 45

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.1, Page 46

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: Tim Rebold Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.2, Page 47

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.2, Page 48

5 x 10 Spiral Out and In of Spacecraft

2 Spiral Out from Earth Spiral In to Moon 1.5

1

0.5

0

(km)

hat -0.5 y -1

-1.5

-2 -2.5

-2 -1 0 1 2 3 4 x (km) 5 hat x 10 Fig. 5.3.2-1 100g Payload Trajectory. (Levi Brown)

Author: Levi Brown Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.2, Page 49

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: Kara Akgulian Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.2, Page 50

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: Levi Brown Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.3, Page 51

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.

2 m

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

Variable Value Units Wet Mass 169 kg Dry Mass 30 kg Required Power 1,540 Watts Burn time 365 days Purchase Cost $1.03 million 2009 dollars

Author: Brad Appel Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.3, Page 52

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

Variable Value Units Thrust 78.5 mN Specific Impulse 1950 s Mass Flow Rate 4.1 mg/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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.3, Page 53

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

Variable Value Units 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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.3, Page 54

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.3, Page 55

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: Brad Appel Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.4, Page 56

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 Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.4, Page 57

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 (퐻2푂2) 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

퐻2푂2 thruster 0.36 1,500 -- 퐻2푂2 propellant 0.02 100 -- Inert Mass 1.9 1,000 -- Totals 19.8 403,000 101.4

Author: Brian Erson Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.4, Page 58

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 퐻2푂2 for its low weight and cost. The integration requires a small increase in inert mass, but not enough to eliminate 퐻2푂2 from contention. The propellant is not currently used on the OTV, but the preceding factors outweigh the possibility of using a different propellant. 퐻2푂2 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

Author: Brittany Waletzko Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.4, Page 59

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: Brittany Waletzko Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 60

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: Joshua Elmshaeuser Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 61

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.

Author: Trenten Muller Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 62

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: Trenten Muller Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 63

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

Author: Michael Christopher Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 64

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.

Author: Michael Christopher Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 65

Fig. 5.3.5-3 Pivot in "normal" position. (Michael Christopher)

Fig. 5.3.5-4 Pivot in 180 degrees from "normal" position. (Michael Christopher)

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.

Author: Michael Christopher Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.5, Page 66

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: Michael Christopher Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.6, Page 67

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.

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.6, Page 68

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

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.7, Page 69

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)

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.7, Page 70

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 105 Watts (during translunar) 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

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.7, Page 71

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 Parameter Value Units 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)

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.7, Page 72

 Power Processing Unit (PPU)  Direct Current / Direct Current (DC/DC) Converters

Not to

scale 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.

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Transfer Section 5.3.7, Page 73

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

Author: Ian Meginnis Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 74

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

Variable Value Units 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 Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 75

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: Ryan Nelson Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 76

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.

Author: Korey Lemond Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 77

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

Author: Joshua Elmshaeuser Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 78

Fig. 5.4.1-2 External View of the Lander with legs deployed. (Joshua Elmshaeuser)

1 meter

H2O2 tank

Author: Joshua Elmshaeuser Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 79

Fig. 5.4.1-2 Close up of the Landers internal components with legs stowed. (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 80

1.3 meters

Fig. 5.4.1-3 View of the Landers engine systems. (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 100g Payload – Lunar Descent Section 5.4.1, Page 81

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 (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 100g Payload – Lunar Descent Section 5.4.2, Page 82

5.4.2 – Mission Operations

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.

Lunar Parking Orbit

Moon Lunar Descent Transfer Orbit

Final Descent

Fig. 5.4.2-1. An overview of the entire lunar descent. Note: Not to scale (John Aitchison)

Author: John Aitchison Mission Configuration – 100g Payload – Lunar Descent Section 5.4.2, Page 83

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.

2 1

3

4

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

Author: John Aitchison Mission Configuration – 100g Payload – Lunar Descent Section 5.4.2, Page 84

Table 5.4.2-2 Descent Propellant Masses Propellant Action Time (s) 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)

Author: John Aitchison Mission Configuration – 100g Payload – Lunar Descent Section 5.4.2, Page 85

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.

Mare Cognitum

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.

Author: Kara Akgulian Mission Configuration – 100g Payload – Lunar Descent Section 5.4.2, Page 86

Fig. 5.4.2-5 Topographical map of the landing site. (NASA/USGS/LPI/ASU)

Author: Kara Akgulian Mission Configuration – 100g Payload – Lunar Descent Section 5.4.3, Page 87

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.

Author: Thaddaeus Halsmer Mission Configuration – 100g Payload – Lunar Descent Section 5.4.3, Page 88

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

Author: Thaddaeus Halsmer Mission Configuration – 100g Payload – Lunar Descent Section 5.4.3, Page 89

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: Thaddaeus Halsmer Mission Configuration – 100g Payload – Lunar Descent Section 5.4.3, Page 90

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 Parameter Value Units 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 Mission Configuration – 100g Payload – Lunar Descent Section 5.4.3, Page 91

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 Parameter Value Units 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 Parameter Value Units

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 Mission Configuration – 100g Payload – Lunar Descent Section 5.4.3, Page 92

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 oC (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: Saad Tanvir Mission Configuration – 100g Payload – Lunar Descent Section 5.4.4, Page 93

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.

Author: Christine Troy Mission Configuration – 100g Payload – Lunar Descent Section 5.4.4, Page 94

Lander Top View

Lander Side view

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

Variable Value Units 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

Author: Christine Troy Mission Configuration – 100g Payload – Lunar Descent Section 5.4.5, Page 95

5.4.5 – Communication

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) 2.48 Power Usage (W) 60.53 Purchase Price (2009 $) 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: Trenten Muller Mission Configuration – 100g Payload – Lunar Descent Section 5.4.5, Page 96

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.

Author: John Dixon Mission Configuration – 100g Payload – Lunar Descent Section 5.4.5, Page 97

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: Michael Christopher Mission Configuration – 100g Payload – Lunar Descent Section 5.4.6, Page 98

5.4.6 – Thermal Control

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

Author: Kelly Leffel Mission Configuration – 100g Payload – Lunar Descent Section 5.4.6, Page 99

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

Author: Kelly Leffel Mission Configuration – 100g Payload – Lunar Descent Section 5.4.6, Page 100

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

Author: Kelly Leffel Mission Configuration – 100g Payload – Lunar Descent Section 5.4.6, Page 101

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 Components Mass (kg) 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: Kelly Leffel Mission Configuration – 100g Payload – Lunar Descent Section 5.4.6, Page 102

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 Mission Configuration – 100g Payload – Lunar Descent Section 5.4.6, Page 103

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: Tony Cofer and Adham Fakhry Mission Configuration – 100g Payload – Lunar Descent Section 5.4.7, Page 104

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.

Table 5.4.7-1 Power breakdown of system on Lander Item Power Units Propulsion 124 Watts Communication 60.53 Watts Attitude 25.4 Watts Total 209.93 Watts 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.

Author: Adham Fahkry Mission Configuration – 100g Payload – Lunar Descent Section 5.4.7, Page 105

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.

Author: Adham Fahkry Mission Configuration – 100g Payload – Lunar Descent Section 5.4.7, Page 106

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

Author: Adham Fahkry Mission Configuration – 100g Payload – Lunar Descent Section 5.4.7, Page 107

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.

Power Conditioning and Distribution Unit (PCDU) 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:

Author: Adham Fahkry Mission Configuration – 100g Payload – Lunar Descent Section 5.4.7, Page 108

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 Component Quantity Units Mass 1.932 kg Dimensions 0.033 x 0.033 x 0.033 m Temperature Range -20 to 60 C Cost $12,000 - Notes: Cost is in 2009 US dollars.

Author: Adham Fahkry Mission Configuration – 100g Payload – Locomotion Section 5.5.1, Page 109

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: Cory Alban Mission Configuration – 100g Payload – Locomotion Section 5.5.1, Page 110

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.

Thermally insulated box

Space Ball SOLIMIDE Foam

Linear shaped charge

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

Author: Ryan Lehto Mission Configuration – 100g Payload – Locomotion Section 5.5.1, Page 111

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.

Outer Shell

Dust Removing Vibrating Motor

Main Axle & Motor Housing 0.25 m

CPU

Transceiver

Camera

100g Payload

Fig. 5.5.1-2 Space Ball internal components. (Ryan Lehto)

Author: Ryan Lehto Mission Configuration – 100g Payload – Locomotion Section 5.5.1, Page 112

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.

Author: Ryan Lehto Mission Configuration – 100g Payload – Locomotion Section 5.5.1, Page 113

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: Kara Akgulian Mission Configuration – 100g Payload – Locomotion Section 5.5.1, Page 114

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: Cory Alban Mission Configuration – 100g Payload – Locomotion Section 5.5.2, Page 115

5.5.2 – Propulsion

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 Mission Configuration – 100g Payload – Locomotion Section 5.5.2, Page 116

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.

Author: Ryan Lehto Mission Configuration – 100g Payload – Locomotion Section 5.5.3, Page 117

5.5.3 – Communications

Communication Hardware/Configuration 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.

Author: Joshua Elmshaeuser Mission Configuration – 100g Payload –Locomotion Section 5.5.3, Page 118

Communication Power Requirements 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.

Author: John Dixon Mission Configuration – 100g Payload –Locomotion Section 5.5.3, Page 119

Communication Timeline (and completion of GLXP requirements) 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

Item Link Directiona Size [MB] Transmission Time [hr] 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

Item Link Directiona Size [MB] Transmission Time [hr] 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

Author: Michael Christopher Mission Configuration – 100g Payload - Locomotion Section 5.5.4, Page 120

5.5.4 – Thermal Control

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: John Dixon Mission Configuration – 100g Payload –Locomotion Section 5.5.5, Page 121

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.

3m mm

23mm 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.

Author: Jeff Knowlton Mission Configuration – 100g Payload –Locomotion Section 5.5.5, Page 122

Battery Distribution CPU Drive Reserve Motors 34% 6%

Camera Trans- 5% mission 55% Fig. 5.5.5-2 Space ball Battery distribution over locomotion mission phase. (Jeff Knowlton)

Author: Jeff Knowlton Mission Configuration – 100g Payload – Locomotion Section 5.5.6, Page 123

5.5.6 – Mission Operations

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 Is Space Ball Solution Probability Capable? 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 .

Table 5.5.6-2 Lunar Debris Hazard Analysis

Debris Size Encounter Is Space Ball Solution Probability Capable? Small (0-.02m) High Yes Roll Over Medium (0.02m-4m) Medium Yes Avoid or Dodge Large (4m+) Very Low Yes Avoid

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.

Author: Cory Alban Mission Configuration – 100g Payload – Locomotion Section 5.5.6, Page 124

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 Mission Configuration – 100g Payload – Locomotion Section 5.5.6, Page 125

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.

Author: Cory Alban Mission Configuration – 100g Payload – Mission Integration Section 5.6, Page 126

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

Author: Korey LeMond Mission Configuration – 100g Payload – Mission Integration Section 5.6, Page 127

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.

Author: Korey LeMond Mission Configuration – 100g Payload – Mission Integration Section 5.6, Page 128

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.

Author: Korey LeMond Mission Configuration – 100g Payload – Mission Integration Section 5.6, Page 129

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.

Author: Korey LeMond Mission Configuration – 100g Payload – Mission Integration Section 5.6, Page 130

Fig. 5.6-3 Final Mission Configuration. (Korey LeMond)

Author: Korey LeMond Mission Configuration – 100g Payload – Mission Integration Section 5.6, Page 131

Fig. 5.6-4 Final Mission Configuration with dimensions. (Korey LeMond)

Author: Korey LeMond Mission Configuration – 100g Payload – Risk Analysis Section 5.7, Page 132

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.

Author: Solomon Westerman Mission Configuration – 100g Payload – Cost Analysis Section 5.8, Page 133

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.

Author: Solomon Westerman Mission Configuration – 10kg Payload Section 6, Page 134

6 – Mission Configuration 10kg Payload

Author: Solomon Westerman Mission Configuration – 10kg Payload – System Overview Section 6.1, Page 135

6.1 – System Overview

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)

Author: Caitlyn McKay Mission Configuration – 10kg Payload – System Overview Section 6.1, Page 136

Fig. 6.1-2 Schematic of the Lunar Lander hopping 500 meters. (Joshua Elmshauser)

Table 6.1-1 Mission timeline Elapsed Time Event ddd:hh:mm -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.

Author: Caitlyn McKay Mission Configuration – 10kg Payload – System Overview Section 6.1, Page 137

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.

Author: Caitlyn McKay Mission Configuration – 10kg Payload – Launch Vehicle Section 6.2.1, Page 138

6.2 – Launch Vehicle

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.

Author: Zarinah Blockton Mission Configuration – 10kg Payload – Launch Vehicle Section 6.2.2, Page 139

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.

Author: Andrew Damon Mission Configuration – 10kg Payload – Launch Vehicle Section 6.2.3, Page 140

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.

Author: Kristopher Ezra Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.1, Page 141

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 Axial Lateral Event Acceleration 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 Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.1, Page 142

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:

퐴푙푙표푤푎푏푙푒 푦푖푒푙푑 푙표푎푑 (표푟 푠푡푟푒푠푠) 푀푆 = − 1 (6.3.1-1) 퐷푒푠푖푔푛 푦푖푒푙푑 푙표푎푑 (표푟 푠푡푟푒푠푠)

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 Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.1, Page 143

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 Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.1, Page 144

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 Components Mass (kg) 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

Author: Tim Rebold Mission Configuration – 10 kg Payload – Lunar Transfer Section 6.3.2, Page 145

6.3.2 – Mission Operations

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.

Author: Levi Brown Mission Configuration – 10 kg Payload – Lunar Transfer Section 6.3.2, Page 146

5 x 10 Spiral Out and In of Spacecraft

2 Spiral Out from Earth Spiral In to Moon 1.5

1

0.5

0

(km) hat

y -0.5

-1

-1.5

-2

-2.5

-1 0 1 2 3 4 x (km) 5 hat x 10 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.

Trajectory Correction Maneuver 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.

Table 6.3.2-1 Correction Maneuver Configuration Parameter Value

Isp (s) 1964 mo (kg) 585.6 Propellant for Correction (kg) 1.5

Author: Levi Brown Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.3, Page 147

6.3.3 – Propulsion on the Orbital Transfer Vehicle

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 Variable Value Units Wet Mass 215 kg Dry Mass 30 kg Required Power 2,043 Watts Burn time 365 days Thrust 104 mN Specific Impulse 1964 s Mass flow Rate 5.4 mg/s

Author: Brad Appel Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.4, Page 148

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 퐻2푂2 thruster hardware needed is the same as the 100g case.

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. 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

퐻2푂2 thruster 0.36 1,500 -- 퐻2푂2 Propellant 1.26 100 -- Inert Mass 3.02 1,000 -- Totals 31.74 403,000 145.4

Author: Brian Erson Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.4, Page 149

Attitude Propellant Use

As mentioned in Section 5.3.4, 퐻2푂2 is an adequate propellant to fulfill all mission requirements.

Space Environment Perturbations The analysis for the 10kg payload is identical to the analysis for the 100g payload detailed in Section 5.3.4.

Author: Brian Erson Mission Configuration – 10g Payload – Lunar Transfer Section 6.3.5, Page 150

6.3.5 – Communication

Communication Hardware/Configuration 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.

Communication Ground Stations 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.

Author: Michael Christopher Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.6, Page 151

6.3.6 – Thermal Control

Electronics Thermal Control 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.

Xenon Tank Thermal Control 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.

Electric Thruster Thermal Control 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 System Component Mass (kg) Dimensions Electronics Ammonia ~ 0 - 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) Xenon Tank Wire Heater 1 - Electric Thruster Aluminum Shroud N/A N/A

Author: Ian Meginnis Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.7, Page 152

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%.

Author: Ian Meginnis Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.7, Page 153

Table 6.3.7-1 10kg Payload OTV Power Budget by Group Group Power Units Propulsion 2029 Watts Communication 0 Watts Attitude 145.4 Watts Power 120 Watts Lunar Lander 105 Watts (during Lunar Transfer) 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) Parameter Value Units Mass 16.89 kg Stowage Volume 0.0858 m3 Deployed Area 8.45 m2

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.

Author: Ian Meginnis Mission Configuration – 10kg Payload – Lunar Transfer Section 6.3.7, Page 154

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 Parameter Value Units Mass (includes housing) 15.93 kg Volume 0.0056 m3 Total Energy 2008 W-hr

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 Component Mass (kg) Volume (m3) Solar Arrays 16.89 0.0858 (stowed) Battery 15.93 0.0056 PCDU 11.49 0.0203 DC/DC Converters 1 0.000218

Author: Ian Meginnis Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 155

6.4 – Lunar Descent

6.4.1 – Structures

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

Author: Joshua Elmshaeuser Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 156

Fig. 6.4.1-1 External View of the Lunar Lander with legs deployed. (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 157

Fig. 6.4.1-2 Close up of the Lunar Lander’s internal components with legs stowed. (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 158

Fig. 6.4.1-3 View of the Lunar Lander’s engine systems. (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 159

Fig. 6.4.1-4 Visualization of the 10kg Lander on the Moon. (Joshua Elmshaeuser)

Author: Joshua Elmshaeuser Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 160

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 Variable Value Units 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 Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 161

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.

Author: Ryan Nelson Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.1, Page 162

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.

Author: Korey LeMond Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.2, Page 163

6.4.2 – Mission Operations

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 Action Time (s) Thrust (N) Radial Burn 000 - 177 1650 Rotation 177 - 183 165 Vertical 183 - 199 1650 Coast to Land 199 - 210 180

Table 6.4.2-2 Descent Propellant Masses Propellant Action Time (s) 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.

Author: Kara Akgulian Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.2, Page 164

Landing Location We chose to land at the same landing site as the 100g payload case. Please refer to Section 5.4.2.

Author: Kara Akgulian Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.3, Page 165

6.4.3 – Propulsion

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)

Performance/Operating Parameters 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.

Author: Thaddaeus Halsmer Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.3, Page 166

Table 6.4.3-1 Hybrid Engine Performance Specifications Parameter Specification Units Thrust Max / Min 1650 / 165 [N] Burn Time 190.4 [s]

Propellant Feed System 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)

Author: Thaddaeus Halsmer Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.3, Page 167

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: Thaddaeus Halsmer Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.3, Page 168

Power Requirements (during operation) 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 hydrogen peroxide tank is sized based on the tank operating pressure and the amount of H2O2 used during the mission. Table 6.4.3-3 describes the pressure, size and mass of the H2O2 tank.

Table 6.4.3-3 Pressure, Size and Mass of the H2O2 tanks Parameter Value Units Operating Pressure 3.07 MPa Volume 0.0867 m3 Thickness 1.26 mm Diameter 0.550 m Mass 3.35 kg

The total volume of the oxidizer tank is the sum of the volume of the H2O2 required for 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).

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 (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 Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.3, Page 169

Table 6.4.3-4 Pressurant System parameters Parameter Value Units Operating Pressure 21 MPa 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

Propulsion System Inert Mass 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 Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.3, Page 170

Table 6.4.3-5 Propulsion system inert mass breakdown Parameter Value Units Main Lander Engine Oxidizer (H2O2) Tank 3.35 kg Pressurant Gas (He) 1.37 kg Pressurant Tank (He) 1.34 kg Nozzle 15.74 kg Combustion Chamber 3.30 kg Injector 0.99 kg Skirts and Bosses 0.33 kg Feed System 0.99 kg Structural Supports 2.74 kg Valves 6.94 kg Hopper Engines Nozzles 6.30 kg Combustion Chambers 0.29 kg Injectors 0.09 kg Skirts and Bosses 0.03 kg Feed Systems 0.09 kg Structural Supports 0.68 kg Valves 1.10 kg Propulsion System Inert Mass 45.67 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: Saad Tanvir Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.4, Page 171

6.4.4 – Attitude Control

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

Author: Josh Lukasak Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.5, Page 172

6.4.5 – Communication

Communication Hardware/Configuration 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.

Communication Power Requirements Please refer to Section 5.4.6 for details regarding the communications power requirements for the Lunar Lander/Hopper.

Communication Timeline (and completion of GLXP requirements) 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.

Author: Trenten Muller Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.6, Page 173

6.4.6 – Thermal Control

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

Author: Kelly Leffel Mission Configuration – 10kg Payload – Lunar Descent Section 6.4.6, Page 174

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 Components Mass (kg) 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.

Author: Kelly Leffel Mission Configuration – 10 kg Payload – Lunar Descent Section 6.4.7, Page 175

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 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 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.

Table 6.4.7-1 Power breakdown of system on Lander Item Power Units Propulsion 148 Watts Communication 54.53 Watts Attitude 25.4 Watts Total 227.93 Watts 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.

Author: Adham Fakhry Mission Configuration – 10 kg Payload – Lunar Descent Section 6.4.7, Page 176

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 We use solar cells to provide enough power to the communication system during the two- 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.

Author: Adham Fakhry Mission Configuration – 10 kg Payload – Lunar Descent Section 6.4.7, Page 177

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

Author: Adham Fakhry Mission Configuration – 10 kg Payload – Lunar Descent Section 6.4.7, Page 178

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.

Table 6.4.7-2 DC-DC converter system Component Quantity Units Mass 0.815 kg Dimensions 0.033 x 0.033 x 0.033 m Temperature Range -50 to 150 C Cost $ 57,500 - Footnote: 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.

Power Conditioning and Distribution Unit (PCDU) 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.

Author: Adham Fakhry Mission Configuration – 10kg Payload – Locomotion Section 6.5.1, Page 179

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.

Author: Alex Whiteman Mission Configuration – 10kg Payload – Locomotion Section 6.5.1, Page 180

200 Moon Hop Trajectory

150

100

50

altitude (m) altitude

0 Touchdown Liftoff

-50

-100 -1000 -900 -800 -700 -600 -500 -400 -300 -200 -100 0 100 range (m)

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 Variable Value Units Time of Flight 134.5 sec Propellant Used 8.6 kg Average Thrust 197 N

Author: Alex Whiteman Mission Configuration – 10kg Payload – Locomotion Section 6.5.2, Page 181

6.5.2 – Propulsion

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)

Performance/Operating Parameters 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.

Author: Thaddaeus Halsmer Mission Configuration – 10kg Payload – Locomotion Section 6.5.2, Page 182

Table 6.5.2-1 Hybrid Locomotion Engine Performance Specifications Parameter Specification Units 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]

Propellant Feed System The locomotion engines use the same oxidizer feed system used for the main engine which was described in Section 6.4.3.

Power Requirements (during operation) The power requirements during the locomotion maneuver remain unchanged from those described in Section 6.4.3.

Author: Thaddaeus Halsmer Mission Configuration – 10kg Payload – Locomotion Section 6.5.3, Page 183

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.

Author: Josh Lukasak Mission Configuration – 10kg Payload – Integration Section 6.6, Page 184

6.6 – Mission Integration

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)

Author: Korey LeMond Mission Configuration – 10kg Payload – Integration Section 6.6, Page 185

Fig. 6.6-2 Final Mission Configuration with dimensions. (Korey LeMond)

Author: Korey LeMond Mission Configuration – 10kg Payload – Risk Analysis Section 6.7, Page 186

6.7 – 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 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:

Table 6.7-1 Vehicle Success Rate

Vehicle Success Rate Dnepr Launch Vehicle 94% Orbital Transfer Vehicle 88% Lunar Lander 88% Hop (redundant) 99% 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 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.

Author: Solomon Westerman Mission Configuration – 10k Payload – Cost Analysis Section 6.8, Page 187

6.8 – Cost Analysis

Table 6.8-1 Total Mission Cost Expense Cost ($M) Dnepr Launch Vehicle 5.77 Orbital Transfer Vehicle 9.52 Lunar Lander / Hopper 5.71 Overhead 8.58 Total Mission Cost 29.6 $/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.

Author: Solomon Westerman Mission Configuration – Large Payload Section 7, Page 188

7 – Mission Configuration Large Payload

Author: Solomon Westerman Mission Configuration – Large Payload – System Overview Section 7.1, Page 189

7.1 – System Overview

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)

Author: Christine Troy Mission Configuration – Large Payload – System Overview Section 7.1, Page 190

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.

Author: Christine Troy Mission Configuration – Large Payload – System Overview Section 7.1, Page 191

Table 7.1-1 Mission Timeline for large payload configuration

Elapsed Time Event (ddd:hh:mm) -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

Author: Christine Troy Mission Configuration – Large Payload – Launch Vehicle Section 7.2.1, Page 192

7.2 – Launch Vehicle

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)

Author: Zarinah Blockton

Mission Configuration – Large Payload – Launch Vehicle Section 7.2.1, Page 193

Table 7.2.1-1 Characteristics of the Falcon 9 Launch Vehicle

Variable Value Units Height 55 m Diameter 3.66 m Lift-off Mass 333,400 kg Mass to 400km orbit 9,953 kg Number of Stages 2 -- Orbit Inclination 28.5 degrees Cost per kilogram to LEO 3,600 $/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)

Author: Zarinah Blockton

Mission Configuration – Large Payload – Launch Vehicle Section 7.2.2, Page 194

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: Andrew Damon

Mission Configuration – Large Payload – Launch Vehicle Section 7.2.3, Page 195

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.

Author: Kris Ezra

Mission Configuration – Large Payload – Lunar Transfer Section 7.3.1, Page 196

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 Mission Configuration – Large Payload – Lunar Transfer Section 7.3.1, Page 197

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 Mission Configuration – Large Payload – Lunar Transfer Section 7.3.1, Page 198

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.

Author: Tim Rebold Mission Configuration – Large Payload – Lunar Transfer Section 7.3.1, Page 199

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.

Author: Korey LeMond Mission Configuration – Large Payload – Lunar Transfer Section 7.3.2, Page 200

7.3.2 – Mission Operations

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: Kara Akgulian Mission Configuration – Large Payload – Lunar Transfer Section 7.3.2, Page 201

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.

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 7.3.2-2.

Table 7.3.2-2 Correction Maneuver Configuration Parameter Value

Isp (s) 2250 mo (kg) 9953 Propellant for Correction (kg) 22.5

Author: Levi Brown Mission Configuration – Large Payload – Lunar Transfer Section 7.3.3, Page 202

7.3.3 – Propulsion on the Orbital Transfer Vehicle

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.

PPU

FCS Thruster

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

Variable Value Units Wet Mass 3,810 kg Dry Mass 520 kg Required Power 38,773 Watts Burn time 365 days Payload Capability 4,545 kg

Author: Brad Appel Mission Configuration – Large Payload – Lunar Transfer Section 7.3.3, Page 203

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

Variable Value Units Thrust 424 mN Specific Impulse 2250 s Mass Flow Rate 19.2 mg/s 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: Brad Appel Mission Configuration – Large Payload – Lunar Transfer Section 7.3.4, Page 204

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 퐻2푂2 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

Our three attitude subsystems: sensors, reaction wheels, and thrusters combine to create a 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 Mission Configuration – Large Payload – Lunar Transfer Section 7.3.4, Page 205

Attitude Propellant Use 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.

퐻2푂2 still fits the bill as a powerful and adequate propellant for the large payload.

Space Environment Perturbations Reaction wheels are used to counteract the attitude perturbations due to the space environment. Refer to Section 5.3.4 for further details.

Author: Brian Erson Mission Configuration – Large Payload – Lunar Transfer Section 7.3.5, Page 206

7.3.5 – Communication

Communication Hardware/Configuration 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.

Communication Ground Stations 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.

Author: Michael Christopher Mission Configuration – Large Payload – Lunar Transfer Section 7.3.6, Page 207

7.3.6 – Thermal Control

Electronics Thermal Control 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.

Xenon Tank Thermal Control For the large payload case, we also employ a simple wire heater to keep the xenon within its storage temperatures.

Electric Thruster Thermal Control 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 System Component Mass (kg) Dimensions Electronics Ammonia ~ 0 - ID = 3.04cm; OD = 3.34cm; Heat Pipes 15.43 Length = 5m Total Cross-Sectional Area = 17.31m2 Radiators 23.4 (8 fins @ 2.163m2 each) Xenon Tank Wire Heater 1 - Electric Thruster Aluminum Shroud N/A N/A

Author: Ian Meginnis Mission Configuration – Large Payload – Lunar Transfer Section 7.3.7, Page 208

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 Propulsion 38773 Watts Communication 0 Watts Attitude 305.4 Watts Power 1731.5 Watts Lunar Lander 105 Watts (during translunar) 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-

Author: Ian Meginnis Mission Configuration – Large Payload – Lunar Transfer Section 7.3.7, Page 209

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) Parameter Value Units Mass 286.4 kg Deployed Area 143.2 m2

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 Parameter Value Units Mass (includes housing) 271.7 kg Volume 0.0957 m3 Total Energy 34261 W-hr

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.

Author: Ian Meginnis Mission Configuration – Large Payload – Lunar Transfer Section 7.3.7, Page 210

Table 7.3.7-4 Large Payload OTV Power Subsystem Dimensions Component Mass (kg) Volume (m3) Solar Arrays 286.4 5.0 (stowed) Battery 271.7 0.0957 PCDU 194.8 0.3437 DC/DC Converters 1 0.000218

Author: Ian Meginnis Mission Configuration – Large Payload – Lunar Descent Section 7.4.1, Page 211

7.4 – Lunar Descent

7.4.1 – Structures

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 100g Large Scale Structural Component Dimension Dimension 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 Mission Configuration – Large Payload – Lunar Descent Section 7.4.1, Page 212

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.

Author: Ryan Nelson Mission Configuration – Large Payload – Lunar Descent Section 7.4.2, Page 213

7.4.2 – Mission Operations

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 Action Time (s) Thrust (N) Radial Burn 000 - 169 34500 Rotation 169 - 174 3450 Vertical 174 - 188 34500

Table 7.4.2-2 Descent Propellant Masses Propellant Action Time (s) 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.

Author: John Aitchison Mission Configuration – Large Payload – Lunar Descent Section 7.4.3, Page 214

7.4.3 – Propulsion

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)

Performance/Operating Parameters 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 Mission Configuration – Large Payload – Lunar Descent Section 7.4.3, Page 215

Table 7.4.3 -1 Hybrid Engine Performance Specifications

Parameter Specification Units Thrust Max / Min 34500 / 3450 [N] Burn Time 188.9 [s]

Propellant Feed System The propellant feed system architecture is unchanged from the system described in Section 5.4.3.

Power Requirements (during operation) 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 hydrogen peroxide tank is sized based on the tank operating pressure and the amount of H2O2 used during the mission. Table 7.4.3-3 describes the pressure, size and mass of the H2O2 tank.

Table 7.4.3-3 Pressure, Size and Mass of the H2O2 tanks Parameter Value Units Operating Pressure 3.07 MPa Volume 1.652 m3 Thickness 3.376 mm Diameter 1.467 m Mass 63.87 kg

Author: Thaddaeus Halsmer/Saad Tanvir Mission Configuration – Large Payload – Lunar Descent Section 7.4.3, Page 216

The total volume of the oxidizer tank is the sum of the volume of the H2O2 required for 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).

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 (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.

Table 7.4.3-4 Pressurant System parameters Parameter Value Units Operating Pressure 21 MPa 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

Propulsion System Inert Mass 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 Mission Configuration – Large Payload – Lunar Descent Section 7.4.3, Page 217

Table 7.4.3-5 Propulsion system inert mass breakdown Parameter Value Units Oxidizer (H2O2) Tank 63.87 kg Pressurant Gas (He) 26.08 kg Pressurant Tank (He) 25.49 kg Nozzle 115.50 kg Combustion Chamber 8.67 kg Injector 2.60 kg Skirts and Bosses 0.87 kg Feed System 2.60 kg Structural Supports 24.57 kg Valves 6.94 kg

Propulsion System Inert Mass 277.18 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.

Author: Saad Tanvir Mission Configuration – Large Payload – Lunar Descent Section 7.4.4, Page 218

7.4.4 – Attitude Control

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.

Author: Josh Lukasak Mission Configuration – Large Payload – Lunar Descent Section 7.4.5, Page 219

7.4.5 – Communication

Communication Hardware/Configuration 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.

Communication Power Requirements Please refer to Section 5.4.6 for details regarding the Communications Power Requirements for the Lunar Lander/Hopper.

Communication Timeline (and completion of GLXP requirements) 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.

Author: Michael Christopher Mission Configuration – Large Payload – Lunar Descent Section 7.4.6, Page 220

7.4.6 – Thermal Control

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

Author: Kelly Leffel Mission Configuration – Large Payload – Lunar Descent Section 7.4.6, Page 221

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

Components Mass (kg) 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.

Author: Kelly Leffel Mission Configuration – Large Payload – Lunar Descent Section 7.4.7, Page 222

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 minutes at peak power consumption. The attitude system uses a peak total power of 25.4 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.

Table 7.4.7-1 Power breakdown of system on Lander Item Power Units Propulsion 275 Watts Communication 54.53 Watts Attitude 25.4 Watts Total 354.93 Watts 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.

Author: Adham Fakhry Mission Configuration – Large Payload – Lunar Descent Section 7.4.7, Page 223

Solar Array Sizing We use solar cells to provide enough power to the communication system during the two- 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.

Author: Adham Fakhry Mission Configuration – Large Payload – Lunar Descent Section 7.4.7, Page 224

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.

Table 7.4.7-2 DC-DC converter system Component Quantity Units Mass 0.985 kg Dimensions 0.033 x 0.033 x 0.033 m Temperature Range -50 to 150 C Cost $ 68,500 - 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.

Power Conditioning and Distribution Unit (PCDU) Please refer to Section 5.4.7 for the complete analysis of why we need a PCDU, what is used for and its specifications.

Author: Adham Fakhry Mission Configuration – Large Payload – Locomotion Section 7.5.1, Page 225

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.

Author: Alex Whiteman Mission Configuration – Large Payload – Locomotion Section 7.5.1, Page 226

150 Moon Hover Trajectory

100

50 altitude (m)

0

-50 -700 -600 -500 -400 -300 -200 -100 0 100 range (m)

Fig. 7.5.1-1 Lander hover trajectory. (Alex Whiteman)

Table 7.5.1-1 Hover Trajectory Results

Variable Value Units Time of Flight 61.4 sec Propellant Used 19.9 kg Average Thrust 1011 N

Author: Alex Whiteman Mission Configuration – Large Payload – Locomotion Section 7.5.2, Page 227

7.5.2 – Propulsion

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.

Performance/Operating Parameters 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

Parameter Specification Units Thrust (average) 3954 [N] Burn Time 61.4 [s]

Propellant Feed System The propellant feed system architecture is unchanged from the system described in Section 7.4.3.

Power Requirements (during operation) During this phase of flight the power requirement is unchanged from what was given in Section 7.4.3.

Author: Thaddaeus Halsmer Mission Configuration – Large Payload – Locomotion Section 7.5.3, Page 228

7.5.3 – Attitude Control

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.

b3

b2

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)

Author: Josh Lukasak Mission Configuration – Large Payload – Vehicle Integration Section 7.6, Page 229

7.6 – Mission Integration

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.

Author: Korey LeMond Mission Configuration – Large Payload – Risk Analysis Section 7.7, Page 230

7.7 – Risk Analysis

The probability of success of the mission is computed by combining the probability of success for each vehicle in the mission. We 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. Table 7.7-1 Vehicle Success Rate

Vehicle Success Rate Falcon 9 Launch Vehicle 99% Orbital Transfer Vehicle 96% Lunar Lander/Hover 97% Mission Success Rate 92%

This mission success rate satisfies the 90% mission requirement in one launch. Our large payload mission can reliably deliver payload to a lunar base.

Author: Solomon Westerman Mission Configuration – Large Payload –Cost Analysis Section 7.8, Page 231

7.8 – Cost Analysis

Table 7.8-1 Total Mission Cost

Expense Cost ($M) Falcon 9 Launch Vehicle 38.2 Orbital Transfer Vehicle 106.6 Lunar Lander / Hover Vehicle 28.6 Overhead 49.1 Total Mission Cost 222.6 $/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.

Author: Solomon Westerman Alternative Designs Section 8.1.1, Page 232

8 – Alternative Designs

Author: Solomon Westerman Alternative Designs Section 8, Page 233

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.

Author: Solomon Westerman Alternative Designs – Launch Vehicle Section 8.1.1, Page 234

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) 1380 13,768 Start (Russia) 632 11,687 Falcon 1e (USA) 1010 9,100 Kosmos 3M(Russia) 1500 8,667 Rockot (Russia) 1850 7,297 Long March 2C (China) 2400 7,031 Strela (Russia) 1700 6,300 Long March 2D (China) 3500 5,000 Dnepr (Russia) 4400 4,800

Table 8.1.1-2 Medium Launch Vehicle Characteristics (Futron Corporation)

Vehicle Payload to LEO-200 km Cost per kg to LEO ($/kg) (kg) Atlas 2 (USA) 8640 11,029 Ariane 44L (ESA) 10200 11,314 Delta 3 (USA) 8291 10,000 Long March 2F (China) 8400 6,700 Long March 2E (China) 9200 5,500 Soyuz (Russia) 7000 5,400 Falcon 9 (USA) 10454 3,600

Author: Zarinah Blockton Alternative Designs – Launch Vehicle Section 8.1.1, Page 235

Table 8.1.1-3 Large Launch Vehicle Characteristics

Vehicle Payload to LLO (kg) Cost per kg to LLO ($/kg) Delta IV M (USA) 6360 26,536 Delta IV M (5,2) (USA) 7705 24,680 Long March 2C (China) 1517 23,568 PSLV (India) 2206 22,470 Delta IV M (5,4) (USA) 9874 20,595 Delta IV M(4,2) (USA) 8900 19,798 Long March 2E (China) 4466 17,288 Delta IV H (USA) 18545 17,251 Atlas V (USA) 7836 16,733 Dnepr (Russia) 2108 12,521 Ares V (USA) 80564 6,788 Falcon 9 (USA) 6780 6,782 Falcon 9 H (USA) 4,973 22234

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.

Author: Zarinah Blockton Alternative Designs – Lunar Transfer Section 8.2.1, Page 236

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 Alternative Designs – Lunar Transfer Section 8.2.1, Page 237

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 Lunar Parking Intermediate Orbit Orbit radius Parameter Hohmann Bielliptic km km km ΔV (km/s) 4.0 4.2 200 110 1 x 106 TOF (days) 5 81

Author: Levi Brown Alternative Designs – Lunar Transfer Section 8.2.1, Page 238

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.

Author: Kara Akgulian Alternative Designs – Lunar Transfer Section 8.2.1, Page 239

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.

ΔV2 ΔV1

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.

Author: Andrew Damon Alternative Designs – Lunar Transfer Section 8.2.1, Page 240

Table 8.2.1-1 Gravitational and Orbital Parameters for the Earth-Moon Hohmann Transfer

Variable Value Units

ΔV1 3.131 km/s ΔV2 0.820 km/s ΔVtot 3.952 km/s ΔVpc 2.166 km/s

Author: Andrew Damon Alternative Designs – Lunar Transfer Section 8.2.1, Page 241

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: Levi Brown Alternative Designs – Lunar Transfer Section 8.2.2, Page 242

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.

Author: Saad Tanvir Alternative Designs – Lunar Transfer Section 8.2.2, Page 243

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)

Author: Kristopher Ezra Alternative Designs – Lunar Transfer Section 8.2.2, Page 244

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: Kristopher Ezra Alternative Designs – Lunar Transfer Section 8.2.2, Page 245

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.

Author: Brad Appel Alternative Designs – Lunar Transfer Section 8.2.3, Page 246

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 6.5 1.22 kg + PPU)

Note: PPU is Power Processing Unit; Cost is about $150,000 for either option

Author: Brittany Waletzko Alternative Designs – Lunar Transfer Section 8.2.3, Page 247

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).

Author: Tony Cofer Alternative Designs – Lunar Transfer Section 8.2.3, Page 248

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.

Author: Jeff Knowlton Alternative Designs – Lunar Transfer Section 8.2.3, Page 249

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

Author: Ian Meginnis Alternative Designs – Lunar Transfer Section 8.2.3, Page 250

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.

Author: Ian Meginnis Alternative Designs – Lunar Descent Section 8.3.1, Page 251

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.

Author: Trenten Muller Alternative Designs – Lunar Descent Section 8.3.1, Page 252

Figure 8.3.1-1 shows the horizontal slide distances of our Lander that were computed.

30

25

20

15 distance (km) 10

5 10g 15g 20g 0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 horizontal velocity (km/sec)

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.

Author: Trenten Muller Alternative Designs – Lunar Descent Section 8.3.1, Page 253

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)

Author: Christine Troy Alternative Designs – Lunar Descent Section 8.3.1, Page 254

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)

Author: Caitlyn McKay Alternative Designs – Lunar Descent Section 8.3.1, Page 255

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.

Author: Caitlyn McKay Alternative Designs – Lunar Descent Section 8.3.1, Page 256

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.

Author: Andrew Damon Alternative Designs – Lunar Descent Section 8.3.1, Page 257

Burst valve

0.3 m

0.5 m

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.

Author: Andrew Damon Alternative Designs – Lunar Descent Section 8.3.1, Page 258

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.

Author: Kristopher Ezra Alternative Designs – Lunar Descent Section 8.3.1, Page 259

5 x 10 Work/Energy Based Accelerations 3

2.5

2

1.5 Accelleration (g) 1

0.5

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5 Collision Distance (m) 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)

Author: Kristopher Ezra Alternative Designs – Lunar Descent Section 8.3.1, Page 260

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.

Author: Alex Whiteman Alternative Designs – Lunar Descent Section 8.3.1, Page 261

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.

w 2v v v=0

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

Author: Kristopher Ezra Alternative Designs – Lunar Descent Section 8.3.1, Page 262

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.

Linear Velocity vs Tether Length 2

1.8

1.6

1.4

1.2

1

0.8

0.6

Magnitude of Linear Velocity of Linear Magnitude (km/s) 0.4

0.2

0 0 10 20 30 40 50 60 Tether Length (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)

Author: Kristopher Ezra Alternative Designs – Lunar Descent Section 8.3.1, Page 263

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.

Author: Kristopher Ezra Alternative Designs – Lunar Descent Section 8.3.1, Page 264

Mass Savings 100

50

0

-50

-100

-150

-200 Mass Savings (kg) -250

-300

-350

-400 0 10 20 30 40 50 60 Tether Length (km)

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)

Author: Kristopher Ezra Alternative Designs – Lunar Descent Section 8.3.1, Page 265

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: Josh Lukasak Alternative Designs – Lunar Descent Section 8.3.2, Page 266

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.

Author: Ryan Nelson Alternative Designs – Lunar Descent Section 8.3.3, Page 267

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 휃, 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.

y

f

휃 g

x 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 휃,

휃0, a, and f. Please see the MATLAB codes LinearTangentSteeringLaw.m as well as LinearTangentSteeringLaw_func.m for the process used to accomplish this.

Author: John Aitchison Alternative Designs – Lunar Descent Section 8.3.3, Page 268

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: John Aitchison Alternative Designs – Lunar Descent Section 8.3.4, Page 269

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.

Author: Saad Tanvir Alternative Designs – Lunar Descent Section 8.3.4, Page 270

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: Thaddaeus Halsmer Alternative Designs – Lunar Descent Section 8.3.4, Page 271

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.

Author: Brad Appel Alternative Designs – Lunar Descent Section 8.3.5, Page 272

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.

Author: Christine Troy Alternative Designs – Lunar Descent Section 8.3.5, Page 273

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) 푀 푚 = (8.3.5 − 1) 푔 ∗ 퐼푠푝 ∗ 퐿 In Eq. 8.3.5-1 푚 is the mass flow rate in kg/s of propellant. Our goal when using this equation is to keep 푚 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

Author: Josh Lukasak Alternative Designs – Lunar Descent Section 8.3.5, Page 274

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.

Author: Josh Lukasak Alternative Designs – Lunar Descent Section 8.3.5, Page 275

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.

Author: Josh Lukasak Alternative Designs – Lunar Descent Section 8.3.6, Page 276

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.

Author: Kelly Leffel Alternative Designs – Lunar Descent Section 8.3.7, Page 277

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.

Author: Kelly Leffel Alternative Designs – Rover Section 8.4, Page 278

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

Author: Caitlyn McKay Alternative Designs – Rover Section 8.4, Page 279

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.

Author: Caitlyn McKay Alternative Designs – Rover Section 8.4.1, Page 280

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.

CPU

Transceiver Camera Battery

Wheel Drive Assemblies

Fig. 8.4.1- 1 The system configuration for the small rover design. (Ryan Lehto)

Author: Ryan Lehto Alternative Designs – Rover Section 8.4.2, Page 281

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: Joshua Elmshaeuser Alternative Designs – Rover Section 8.4.3, Page 282

8.4.3 – Propulsion

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.

Motor Gear Wheel 0.076 m

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

Author: Ryan Lehto Alternative Designs – Rover Section 8.4.3, Page 283

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: Ryan Lehto Alternative Designs – Rover Section 8.4.4, Page 284

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.

Author: John Dixon Alternative Designs – Rover Section 8.4.5, Page 285

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.

Author: Jeff Knowlton Alternative Designs – Rover Section 8.4.6, Page 286

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).

Copper Lining

Explosive Charge

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.

Author: Caitlyn McKay Alternative Designs – Rover Section 8.4.6, Page 287

Motor Support Beam

Platform m

Steel Cable 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.

Author: Caitlyn McKay Alternative Designs – Rover Section 8.4.7, Page 288

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.

Author: Caitlyn McKay Alternative Designs – Other Locomotion Section 8.5.1, Page 289

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

Author: Ryan Lehto Alternative Designs – Other Locomotion Section 8.5.2, Page 290

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.

Author: Zarinah Blockton Alternative Designs – Other Locomotion Section 8.5.2, Page 291

Table 8.5.2-1 Spring Launch System Results

Variable 100 gram 10 kilogram Units Mass 7.6 18.5 kg Free Length 0.1524 0.1524 m Spring Constant 18879 18879 N/m Velocity 41.6 40.7 m/s Number of Springs 30 70 -- 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 540 518 m Launch g‟s 3549 3402 -- Landing g‟s 26 25.5 -- Spring length for 30 g 18 17 m launch

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: Zarinah Blockton Alternative Designs – Other Locomotion Section 8.5.3, Page 292

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.

Author: Cory Alban About The Authors Section 9, Page 293

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

Author: Solomon Westerman About The Authors Section 9, Page 294

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

Author: Solomon Westerman About The Authors Section 9, Page 295

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

Author: Solomon Westerman About The Authors Section 9, Page 296

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.

Author: Solomon Westerman About The Authors Section 9, Page 297

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.

Author: Solomon Westerman About The Authors Section 9, Page 298

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

Author: Solomon Westerman About The Authors Section 9, Page 299

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.

Author: Solomon Westerman References Section 10, Page 300

10 – References

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