Project Alshlin - University Of Maryland · 2009-09-30 · Project Alshlin: A Lunar Flying yehicle for Rapid Universal Surface Access ENAE484 2009 Class Final Project Submission Dr

Project Alshain: A Lunar Flying Vehicle for Rapid

Universal Surface Access

ENAE484 2009 Class Final Project SubmissionDr. David Akin

Dr. Mary Bowden

ENAE484 Class:Report Editors: Michael Sotak (Compiler and Editor), Alex Janas, Nick D’Amore, Andrew

Turner, Bree McNerney, Nitin Sydney

AvionicsNate NilesNick D'AmoreKush PatelJolyon Zook

Crew SystemsAndrew TurnerPratik DavéBree McNerney

Systems IntegrationAlex JanasNeal VasilakRyan LeboisTheodor TalvacMichael Sotak

Mission PlanningSarah BealAndrew WilsonAdam KirkFazle SiddiqueZach Neumann

Power Propulsion and ThermalAmirhadi EkramiNitin SydneyScott WeinbergArber MasatiMatt Kosmer

Loads Structures And MechanismsAndrew McLarenAdam HalperinEdwin FernandesJoseph ParkBryan HanJarred Alexander Young

Contents

1 Introduction 111.1 Constellation Program (Mike Sotak) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

1.1.1 Altair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111.1.2 Lunar Base . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

1.2 Project Overview (Mike Sotak) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 121.2.1 Potential for Lunar Flying Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 121.2.2 Project Goals and Mission Pro�le . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

1.3 Level One Requirements (Ryan Lebois) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2 Alshain 142.1 Overview of Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

2.1.1 Vehicle Con�guration (Alex Janas) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142.2 Systems Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

2.2.1 Mission Planning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.2.2 Avionics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.2.3 Crew Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.2.4 Loads, Structures, and Mechanics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.2.5 Power, Propulsion, and Thermal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.2.6 Systems Integration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

2.3 Egress from Altair Lander (Sarah Beal) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.3.1 Resources Onboard Altair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172.3.2 Mission Plan for Egress from Altair . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

3 Mission Planning 213.1 Range Versus Vehicle Mass (Adam Kirk/ Neal Vasilak) . . . . . . . . . . . . . . . . . . . . . 21

3.1.1 Purpose . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213.1.2 Assumptions and Calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213.1.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

3.2 Exploration Range (Sarah Beal) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233.2.1 Solar Particle Event . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233.2.2 Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

3.3 Area Surrounding Altair Landing Site . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 253.3.1 Local Terrain of Lunar Base Site (Adam Kirk) . . . . . . . . . . . . . . . . . . . . . . 253.3.2 Crater Analysis (Zach Neumann) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

3.4 Payload Bay Sizing (Fazle Siddique) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 323.5 Locking Mechanism for Payload Bay (Sarah Beal) . . . . . . . . . . . . . . . . . . . . . . . . 343.6 Refueling Alshain (Andrew Wilson) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 343.7 Dust Maintenance (Adam Kirk) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

4 Propulsion System 364.1 System Overview (Nitin Sydney) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 364.2 Main Engine System (Matt Kosmer) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

4.2.1 Thrust Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 374.2.2 Number of Engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 384.2.3 Selection of Performance Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . 39

4.3 Propulsion Analysis (Nitin Sydney) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 424.3.1 Analysis Requirments and Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . 424.3.2 Propellant Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.4 Tank System (Nitin Sydney) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444.4.1 Tank Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444.4.2 Propellant Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444.4.3 Pressurant Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 474.4.4 Tank System Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

2

4.5 Valves, Pressure Lines and Pressure Regulators (Nitin Sydney) . . . . . . . . . . . . . . . . . 504.5.1 Valves and Pressure Regulators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 504.5.2 Pressure and Propellant Lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

4.6 Reaction Control System (Scott Weinberg) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524.6.1 Control Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524.6.2 Thruster Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 534.6.3 System Con�guration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 584.6.4 Two-Fault Tolerance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 604.6.5 Piping and Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 614.6.6 Summary of Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

5 Thermal (Amirhadi Ekrami) 625.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 625.2 Lunar Environmental Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 635.3 Seating Structure Heat Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 635.4 Electronics and Fuel Cells Heat Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

5.4.1 Electronics Cooling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 645.4.2 Electronics Heating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 655.4.3 Fuel Cell Heat Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

5.5 Cryogenic Tank Heat Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 675.6 Equator Design Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

6 Structures and Mechanics 706.1 Coordinate System (Edwin Fernandes) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 706.2 Center of Gravity Analysis (Jarred Young) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 706.3 Moment of Inertia Analysis (Jarred Young) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 736.4 Landing Gear (Andrew McLaren) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

6.4.1 Number of Legs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 796.4.2 Landing Load Acceleration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 796.4.3 Landing Tipping Acceleration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 806.4.4 Ratcheting Spring Lock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 816.4.5 Cold Rated Spring (Bryan Han) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 816.4.6 Leg Interface to Main Structure (Adam Halperin) . . . . . . . . . . . . . . . . . . . . 816.4.7 Leg Structural Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 816.4.8 Leg Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 836.4.9 Thermal Shielding (Bryan Han) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83

6.5 Truss Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 836.5.1 Design Rationale (Adam Halperin) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 836.5.2 Structure Inventory (Bryan Han) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84

6.6 Crew Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 856.6.1 Platform Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 856.6.2 Roll Cage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86

6.7 Support Structure (Adam Halperin) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 866.7.1 Support Structure Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 866.7.2 Support Structure Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 866.7.3 Support Structure Shape . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 876.7.4 Beam Choice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 876.7.5 I-beam Height Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 886.7.6 I-Beam Thickness Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 896.7.7 Tubular Beam Thickness Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 896.7.8 Support Beam Loading Cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 906.7.9 Support Structure Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 916.7.10 Truss System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91

6.8 Vibration Analysis (Edwin Fernades, Joe Park) . . . . . . . . . . . . . . . . . . . . . . . . . . 97

3

6.8.1 Random Vibration Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 986.9 Roll Cage (Jarred Young) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100

7 Crew Systems 1037.1 Crew Placement Con�guration (Breanne McNerney) . . . . . . . . . . . . . . . . . . . . . . . 103

7.1.1 Seating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1037.1.2 Sightlines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1057.1.3 Restraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105

7.2 Crew Interface (Andrew Turner) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1087.2.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1087.2.2 Joysticks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1087.2.3 Controls Input . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1097.2.4 Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1097.2.5 Control Panel Layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1107.2.6 Lighting (Pratik Davé) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111

7.3 Debris Protection (Pratik Davé) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1147.3.1 Protection from lunar meteoroids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1147.3.2 Protection during takeo� and landing . . . . . . . . . . . . . . . . . . . . . . . . . . . 116

7.4 Contingency Procedures (Andrew Turner) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1207.4.1 Extra Consumables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1207.4.2 Radiation Contingency (Breanne McNerney) . . . . . . . . . . . . . . . . . . . . . . . 122

8 Hardware (Beal, Vasilak, McNerney, D'Amore, Sotak, Turner) 1238.1 Egress/Ingress (Sarah Beal) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123

8.1.1 Winch (Sarah Beal) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1238.1.2 Elevator (Neal Vasilak) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1248.1.3 Chosen Design (Neal Vasilak) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 126

8.2 Crew Stations (Breanne McNerney) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1268.2.1 Ingress/Egress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1268.2.2 Restraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1278.2.3 Incapacitated Astronaut . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127

9 Avionics 1289.1 General Overview (Kush Patel) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128

9.1.1 Requirements (Kush Patel and Nick D'Amore) . . . . . . . . . . . . . . . . . . . . . . 1289.2 External dependencies (Kush Patel) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129

9.2.1 NASA Lunar Surface Architecture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1299.3 Communications (Jolyon Zook) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 130

9.3.1 Frequencies Speci�cations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1309.3.2 Architectural Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1319.3.3 Antennas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 131

9.4 Command and Data Handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1329.4.1 Equipment (Kush Patel) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132

9.5 Guidance (Nick D'Amore) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1339.5.1 Overview of Flight Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1339.5.2 Landing Site Selection and Hazard Avoidance . . . . . . . . . . . . . . . . . . . . . . . 1339.5.3 Control Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134

9.6 Navigation (Nate Niles) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1359.6.1 Position Estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1359.6.2 Attitude Estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1369.6.3 Flight Path Propagation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1379.6.4 Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1379.6.5 Error budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 138

9.7 Vehicle Status Monitoring and Fault Handling (Nick D'Amore) . . . . . . . . . . . . . . . . . 141

4

9.7.1 Fault Tolerance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1419.8 Control (Nate Niles) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142

10 Power (Arber Masati) 14310.1 Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 144

10.1.1 In-Flight Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14410.1.2 Landed Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14410.1.3 Contingency Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14510.1.4 Summary of Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 146

10.2 Li-ion Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14610.2.1 Rechargeable Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14610.2.2 Non-Rechargeable Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 147

10.3 Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14810.3.1 PEM Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14810.3.2 Fuel Cell Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14910.3.3 Reactants Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15010.3.4 Water Management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 152

10.4 Contingency Options . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15210.5 Power Management and Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15210.6 Wiring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15210.7 Summary of Power System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153

11 Design Reference Mission (Andrew Wilson) 15411.1 Mission Site . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15411.2 Payloads Carried . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15411.3 Delta V requirements (Adam Kirk) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 157

11.3.1 Assumptions/Calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15711.3.2 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15711.3.3 Reference to MATLAB scripts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 157

12 Conclusions 15812.1 Costing Analysis (Theo Talvac) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 158

12.1.1 NASA Cost Models . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15812.1.2 In�ation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15812.1.3 Spacecraft/Vehicle Level Cost Model (SVLC Model) . . . . . . . . . . . . . . . . . . . 15812.1.4 Advanced Missions Cost Model (AMC Model) . . . . . . . . . . . . . . . . . . . . . . 16012.1.5 Preliminary Estimate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161

12.2 Reliability Fault Tree (Ryan Lebois) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16112.2.1 Parts List . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16212.2.2 Fault Tree Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16212.2.3 Subsystems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16212.2.4 Monte Carlo Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16412.2.5 Results: Loss of Crew . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16412.2.6 Results: Loss of Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164

12.3 Technology Readiness Levels (Theo Talvac) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16512.3.1 De�nition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16512.3.2 TRL List (Theo Talvac, Mike Sotak) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 165

12.4 Outreach (Ryan Lebois) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16512.4.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16712.4.2 Maryland Day . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16712.4.3 University Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16712.4.4 Community Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16812.4.5 Additional Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17012.4.6 Outreach Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 170

5

12.5 Mass Budget (Neal Vasilak) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17012.5.1 Total Inert Mass Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17012.5.2 Percentage Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171

A Appendix 173A.1 Trade Studies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173

A.1.1 One vs Two Person Vehicle (Neal Vasilak) . . . . . . . . . . . . . . . . . . . . . . . . . 173A.1.2 Battery Mass Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173A.1.3 Con�guration Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175A.1.4 Materials Study (Jarred Young) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 188A.1.5 Moment of Inertia Analysis (Jarred Young) . . . . . . . . . . . . . . . . . . . . . . . . 189A.1.6 Delta V Requirements for Multi-Crater Missions (Adam Kirk) . . . . . . . . . . . . . 193A.1.7 Glide vs Hop Trade Study (Adam Kirk) . . . . . . . . . . . . . . . . . . . . . . . . . . 194

A.2 MATLAB Scripts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195A.2.1 Delta V Analysis (Adam Kirk) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195A.2.2 Multi-Crater Analysis (Adam Kirk) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 197A.2.3 Range vs. Vehicle Mass Trade study (Adam Kirk, Neal Vasilak) . . . . . . . . . . . . 198A.2.4 Tank Sizing Scripts (Nitin Sydney) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200A.2.5 Thermal Scripts (Amirhadi Ekrami) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203A.2.6 Moment of Inertia Scripts (Jarred Young) . . . . . . . . . . . . . . . . . . . . . . . . . 203A.2.7 Payload Bay Sizing (Fazle Siddique) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

A.3 Payload Database (Fazle Siddique) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206A.4 Ballistic Hop and Glide Equation Derivations (Adam Kirk) . . . . . . . . . . . . . . . . . . . 209

A.4.1 Ballistic Hop on Flat Airless Body . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209A.4.2 Ballistic Hop on Flat Airless Body with Elevation Change . . . . . . . . . . . . . . . . 210A.4.3 Propulsive Glide on Flat Airless Body . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

A.5 Dummy Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212A.6 Outreach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212

A.6.1 Hours List . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212A.7 Additional References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213

List of Figures

1 Moon Base . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112 Initial Design Concept from PDR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153 Con�gurations 2 and 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164 Con�gurations 4 and 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 165 Final Con�guration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 176 Altair lander with Tri-ATHLETE and Power Supply Unit stowed on the cargo deck. . . . . . 187 Tri-ATHLETE Con�gurations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 188 Power Supply Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199 Lunar Surface Manipulation System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1910 Crew Mobility Chassis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1911 Distances from the base site and the required ΔV . . . . . . . . . . . . . . . . . . . . . . . . 2112 Mass vs Max Range Trade Study Plot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2213 Plot of the relationship between astronauts food speed and stride length. . . . . . . . . . . . 2314 Exploration range of male and female astronauts as a function of time of �ight. . . . . . . . . 2415 South Pole elevation map around lunar base site . . . . . . . . . . . . . . . . . . . . . . . . . 2516 South Pole slope map around lunar base site . . . . . . . . . . . . . . . . . . . . . . . . . . . 2617 South Pole % sunlight map around lunar base site . . . . . . . . . . . . . . . . . . . . . . . . 2618 Shackleton Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2819 Amundsen Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2820 Cabeus Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

6

21 De Gerlache Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2922 Faustini Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3023 Malapert Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3024 Nobile Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3125 Scott Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3126 Shoemaker Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3227 Wiechert Crater . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3228 Payload Bay Mass Plot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3329 Payload Bay Volume Plot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3430 Propulsion System Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3631 System Mass vs Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3832 Expansion Ratio vs ISP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4133 Engine Diameter vs Engine Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4234 Propellant Tank Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4635 Mass Minimzation Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4736 Total Tank Mass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4937 Valve and PR System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5038 Mechanical Pressure Regulator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5139 Burn Time vs. Mass of Thrusters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5540 Thrust vs. Mass Flow for 6 and 8 thruster clusters . . . . . . . . . . . . . . . . . . . . . . . . 5641 Chamber Pressure vs. Nozzle length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5742 Chamber Pressure vs. Nozzle Diameter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5743 RCS Con�guration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5844 Boom Truss . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6045 Optical Solar Re�ector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6446 Radiation Area Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6547 Thermal Louvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6648 Multi Layer Insulation Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6849 Passive Radiation Comparison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6850 Tilt Angle E�ect . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6951 Coordinate System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7052 Coordinate System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7053 Alshain Side Image . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7254 Sample Excel CG Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7555 Number of Legs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7956 Landing Tipping Acceleration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8057 Leg Structure Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8258 Design Rationale Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8459 Support Structure Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8660 Support Structure Shape . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8761 I-Beam vs Tube Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8862 I-Beam Height Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8863 IBeam Thickness Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8964 Support Beam Loading Cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9065 Beam Support Shear and Moment Diagrams . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9166 Tubular Member Mass vs Radius . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9367 Support Structure Crossbeams . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9368 Fuel Tank Support Side View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9569 Fuel Tank Support Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9570 Fuel Support Structure Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9671 Full Skeleton Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9772 First Mode Displacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9873 First Mode Displacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9874 Alshain with Roll Cage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100

7

75 Circular Arch Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10176 Crew Seating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10377 Seating Dimensions Front View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10478 Seating Dimensions Side View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10579 Helmet Sightlines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10680 Lower Body Sightlines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10681 Forward Crew Member Sightline Obstruction . . . . . . . . . . . . . . . . . . . . . . . . . . . 10682 PLSS Restraint Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10783 Boot Restraint Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10784 Left Joystick . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10885 Right Joystick . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10886 Keypad . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10987 Voice Commands (Courtesy newscientist.com) . . . . . . . . . . . . . . . . . . . . . . . . . . . 11088 HUD (Courtesy Dr. William J. Clancy Desert-RATS Mission Simulation, 2006) . . . . . . . . 11089 Warning Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11190 NASA Astronaut Controls Test (Courtesy JSC2000E03273 Fig. 14) . . . . . . . . . . . . . . . 11191 Hardware Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11292 Vehicle lighting con�gurations for (a) the crew �ight control area, (b) the crew ingress/egress

area, and (c) the cargo elevator area. One 20 W halogen lamp is located at each L. . . . . . 11393 Surround Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11494 Meteoroid Flux as a Function of Size . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11695 Conceptual image of TMG body debris shield . . . . . . . . . . . . . . . . . . . . . . . . . . . 11796 Conceptual image of Polycarbonate head debris shield . . . . . . . . . . . . . . . . . . . . . . 11997 Human Needs (Courtesy Dr. David Akin's ENAE484 Life Support Lecture) . . . . . . . . . . 12098 Winch Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12399 Elevator Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124100 Elevator Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125101 Elevator Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 126102 Hardware Ingress Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127103 Hardware Ingress Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127104 Hardware Ingress Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128105 PLSS Restraints Hardware Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128106 Incapacitated Astronaut Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129107 Control System Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143108 Power Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143109 A rate of nC corresponds to a full discharge in 1/n h (Courtesy Kang, B. and Ceder, G.

�Battery materials for ultrafast charging and discharging� Nature Vol. 458. March 12th 2009:pg. 190-193) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 147

110 Fuel Cell (Courtesy http://www.fuelcells.org/basics/how.html) . . . . . . . . . . . . . . . . . 148111 Fuel Cell Sizing Graph (Courtesy Barber, Frano. PEM Fuel Cells Theory and Practice.

Elsevier Academic Press, 2005. pg 58) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150112 Flight Path to Shackleton Crater Basin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154113 Ballistic Hop Trajectories for Mission to Shackleton Crater . . . . . . . . . . . . . . . . . . . 155114 Quadrupole Mass Spectrometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 155115 VAPoR mass spectrometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 156116 Mars Underground Mole . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 156117 Fault tree structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163118 Showing o� the Alshain Mock-up at Maryland Day . . . . . . . . . . . . . . . . . . . . . . . . 168119 Listening to students' reports at one of the Maryland Engineering Challenges . . . . . . . . . 169120 Measuring range of �ight at on of the Maryland Engineering Challenges . . . . . . . . . . . . 169121 Outreach Hours List . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213

8

List of Tables

1 Coordinate Data for Related Sites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 272 Gravity Drag Penalty . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 433 Max Mission Masses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444 Propellant Tank Material Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 455 One Spherical Tank Mass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 456 Spherical Tank Volumes and Masses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 467 Tank Mass Study Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 478 Pressurant Choice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 499 Tank Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4910 Tank Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5011 Valve and Pressure Regulator Estimations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5112 Pressure and Propellant Lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5213 Engine thrust and g-loads for varying burn times . . . . . . . . . . . . . . . . . . . . . . . . . 5414 Six engine thrust with equivalent 8 engine mass for varying burn times/ g-loads . . . . . . . . 5415 Thruster Speci�cations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5816 Summary of RCS Thrust and Moment Control . . . . . . . . . . . . . . . . . . . . . . . . . . 5917 Roll Rates (Angular Acceleration) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5918 Boom Truss Structural Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6019 Mass Breakdown for RCS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6220 Operating Temperature Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6221 Lunar Environment Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6322 Avionics Box Electronics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6423 Emergency Avionics Box Electronics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6624 Cryogenic Propellant Properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6725 Summary of CG Cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7126 Alshain Mass Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7227 Sample Excel CG Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7428 CG Location Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7629 Max CG Shift Per Dimension . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7630 Moments of Inertia Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7831 FIB Info . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8332 Tube Inventory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8433 Margin of Safety Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8534 Support Structure Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9135 Support Structure Margin of Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9236 Random Vibration Loading X-Y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9837 Random Vibration Loading Z . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9938 Damping Ratio vs Frequency . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9939 Quasi-Static Vibration Loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9940 Roll Cage Component Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10141 Rover Seat Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10442 Lighting Con�guration Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11543 Breakdown of TMG debris shield layers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11744 Material Comparison - PMMA (Acrylic) vs. Polycarbonate . . . . . . . . . . . . . . . . . . . 11845 Summary of debris protection design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11946 Drinking Water Tank Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12047 Coolant Water Tank Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12148 Box Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12149 Avionics Mass Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13350 Avionics Volume Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13351 Navigation Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13752 Navigation system uncertainties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 138

9

53 Sample Uncertainty Propagations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14154 Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14455 Landed Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14556 Contingency Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14557 Summary of Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14658 Non-Rechargeable Batteries (Courtesy http://www.quallion.com/sub-tc-primary.asp) . . . . . 14859 PEM Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14960 Reactants Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15161 Reactants Supply Approximation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15162 Wiring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15363 Power System Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15364 Costing Input . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15965 Costing Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15966 SVLC Development Costs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15967 AMC Input Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16068 2004 Costs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16069 AMC Total Costs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16170 AMC Development Costs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16171 Preliminary Estimate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16172 Cost Readiness Levels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16273 Component Reliabilities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16374 Loss of Crew Weak Points . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16475 Loss of Mission Weak Points . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16476 TRL De�nitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16577 TRL List . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16678 Outreach Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17079 Total Inert Mass Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171

10

1 Introduction

1.1 Constellation Program (Mike Sotak)

Since the Apollo program, the United States has foregone going to the Moon in order tofocus on other space applications, including Mars exploration and satellite communications.However, with the retirement of the Space Shuttle Program, NASA has plans to develop theConstellation Program, which will involve not only a manned return to the moon, but willalso lead to the establishment of a Lunar base near the south pole of the Moon. This willpotentially lead to major exploration and research projects on the Moon.

1.1.1 Altair

The Altair lander is the lunar landing module of the Constellation Program. It is the meansfor delivering both crew and cargo to the surface of the Moon. Altair's speci�cations includea cargo deck measuring approximately six meters in diameter and uses liquid oxygen (LOX)and liquid hydrogen (LH2) propellants as a means of propulsion.

1.1.2 Lunar Base

A base on the Moon requires several assumptions, including the development of researchcenters, habitats, and resource mining. It can be assumed that such a base could mine in situmaterials in order to provide the means to establish a long term base. Using Lunar RelaySatellites, astronauts on the Moon should be able to communicate with one another as well aspersons on Earth. Below is an artist's rendition of what a possible moon base might look like:

Figure 1: Moon Base

With such a Moon base, a transportation infrastructure must be developed in order to ef-�ciently travel, research, and explore on the Moon's surface. Since Moon bases have neverbeen developed in the past, the Antarctic base infrastructure was considered as an analogueto what modes of transportation are useful to the exploration of inhospitable areas. For ex-ample, in Antarctica, scientists have the means to travel short distances between buildings

11

and around the base using snowmobiles, the ability to conduct longer, research missions usingclosed cabin vehicles, and even the method of surpassing longer distances quickly using variousaircraft. Similarly, unpressurized rovers could be the �snowmobiles� of the lunar base whilepressurized rovers could conduct the longer, more research oriented missions. However, themeans to travel long distances quickly and reach inaccessible areas is an area that has yet tobe explored in-depth. Thus, a lunar �ying vehicle has been proposed to accomplish these tasksand supplement the South Pole base transportation infrastructure.

1.2 Project Overview (Mike Sotak)

1.2.1 Potential for Lunar Flying Vehicle

A lunar �ying vehicle provides extraordinary potential as a means of transportation on theMoon. Such bene�ts include access to sites otherwise inaccessible to a lunar rover, includingrilles, craters, mountains, and potential lava tubes. To cite a past NASA example, Apollo 15landed next to Hadley Rille during its mission, but had only a rover as a means of transporta-tion. This severely limited the exploration potential of the mission, which could have beengreatly enhanced by a �ying vehicle. Flying vehicles also provide faster transit, thus increasingthe portion of the mission that can be devoted to mission goals rather than transportation. Interms of contingency, lunar �ying vehicles also provide a means of reaching a disabled vehicle,such as a rover, in order to perform crew rescuing operations.

1.2.2 Project Goals and Mission Pro�le

Several goals were developed at the onset of the project in order to provide some level oforganization for the group as a whole. These goals were:

� To design a lunar �ying vehicle to be used in conjunction with the Altair lander, following a speci�cset of level one requirements

� To construct a relevant hardware component to test aspects of the design that cannot be modeledusing engineering software

� To provide outreach to the community to educate and inform the public about NASA and otheraerospace engineering endeavors

1.3 Level One Requirements (Ryan Lebois)

The design of the Alshain lunar �ying vehicle (LFV) was guided by the level one program requirements.These requirements list explicitly what vehicle performance, features, and functionality are necessary. Theywere given as follows:

Mission Planning

� The LFV, along with all necessary support infrastructure, shall be designed to launch on an Altairlander on a cargo delivery mission

� The LFV shall be capable of takeo� and landing on unimproved sites equivalent to those selected forJ-class Apollo missions

� The LFV shall be capable of being o�-loaded from the lander with minimal crew involvement

� The LFV shall be capable of autonomous �ight and landing at a planned base landing site

12

Crew Systems

� LFV shall accommodate crew sized ranging from 95th percentile American male to 5th percentileAmerican woman

� All crew interfaces shall be in accordance with NASA STD-3000

� LFV shall provide contingency life support for nominal mission plus 24 hours contingency

� LFV shall be designed for crew-guided stable �ight in all mission phases

� LFV design shall ensure su�cient direct sight lines and illumination to allow safe �ight in daylight ornight conditions

� LFV shall be designed for simple interfacing for consumables re-supply, maintenance, and servicing

Performance

� LFV shall be designed for a safe landing on a 15° slope

� LFV shall be designed for a worst-case landing gear on top of a 30-cm obstacle

� LFV shall be designed for a maximum safe landing velocity of 3 m/sec vertical, 1 m/sec lateral

Avionics

� LFV shall be capable of being controlled directly, in teleoperation, and autonomously

� LFV shall be capable of communicating at HDTV rates direct to Earth

� LFV shall provide voice/data/video to and from pressure suits during EVAs

� All critical systems shall be two-fault tolerant, with instrumentation for status monitoring

Management

� Design team shall establish and maintain a canonical reference con�guration for all systems, includingbudgets with margins for mass, cost, and power

� Unless otherwise noted, all systems shall be designed in accordance with NASA standards and require-ments

13

2 Alshain

2.1 Overview of Vehicle

2.1.1 Vehicle Con�guration (Alex Janas)

The con�guration selected for Alshain is a unique design that carefully considers each of thevehicle's many subsystems. With little consideration given to lunar �yers since the Apollo era,there are few if any applicable designs that can be used as a reference when determining vehiclelayout. Because applicable heritage designs are not readily available for use as a reference, awide array of con�gurations were considered prior to the �nal selection.

Each potential con�guration was required to ful�ll several key requirements. The most im-portant considerations are listed as follows:

1. Vehicle center of gravity (C.G.) location for each �ight condition.

2. NASA STD-3000, NASA CxP 70024

3. Tipping moment (vertical C.G.)

4. Crew sight lines

5. Contingency condition: loading 2 incapacitated astronauts

6. Ingress/egress accessibility

7. Stowed dimensions aboard Altair

8. Minimize inert mass

Each consideration a�ects a di�erent aspect of the con�guration, but there are many over-lapping and con�icting areas which make achievement of an optimal con�guration di�cult.Although a separate list of level one requirements exists to satisfy design constraints, this listpertains to the considerations given when designing the vehicle layout.

The relative C.G. shift of the vehicle is essential for conserving propellant and maintainingvehicle stability in �ight. This parameter is dictated by the location of crew, cargo, componentsof constant mass (e.g. avionics), and components of variable mass (e.g. propellant). Tocalculate the relative shift, an Excel spreadsheet was created using a point mass approximationof equipment. By entering the mass and center of gravity of each object, the script ran through18 possible loading scenarios and returned the farthest C.G. shift in each direction. Theanalysis and results pertaining to Alshain's con�guration will be discussed in much greaterdetail in section ??(REFERENCE LSM SECTION).

The NASA standard documents dictated many aspects of the design, but most importantly arethe restrictions they placed on crew interface and accessibility. For example NASA CxP70024,the NASA standard for Constellation architecture, states that astronauts must be capable ofviewing each other's mission critical control functions. This dictated the placement of the pilotseat at the fore position in Alshain's �nal con�guration, and eliminated several other possiblecon�gurations.

The vertical location of the C.G. as calculated through the Excel spreadsheet is an essentialparameter to consider when designing the landing gear of the vehicle. A lower center of gravitydecreases the vehicle's tipping moment and thus allows for smaller, more lightweight landinggear.

14

Figure 2: Initial Design Concept from PDR

Crew sight lines are of extreme importance when operating the vehicle under manual control.The crew must be able to have clear vision while in �ight and most importantly be able toavoid obstacles and re�ne their attitude while landing.

The contingency consideration is inferred from the level 1 requirement dictating 24 hoursof contingency. Consideration �ve states that the lunar �ying vehicle shall be capable ofcarrying two incapacitated astronauts while allowing for a third astronaut to pilot the vehicle.This consideration dictated size and ease of access for the payload bay. The ingress/egressconsideration refers to conventional vehicle operations, and each astronaut's ability to get onand o� of the vehicle. Vehicle height and crew operation positions were analyzed in order toful�ll this constraint.

The stowed vehicle dimensions consideration refers to minimizing the volume Alshain willconsume when loaded aboard an Altair lunar lander. Although the Altair and its payloaddeck is still being designed, by minimizing the area Alshain can increase the likelihood that itwill be accommodated by the lander and also leave room for other valuable cargo.

Finally, by minimizing the vehicle's inert mass, it is possible to decrease both the cost oftransport to the lunar surface, and to limit the amount of fuel required for the physicaloperation of the vehicle.

The following images present �ve other con�gurations that were considered when designingAlshain. Each con�guration below was considered for its desirable results for individual vehicleconsiderations, but ultimately exhibited shortfalls either too great or numerous to allow itsuse. The �nal image is that of Alshain's reference con�guration.

15

(a) Con�guration 1 (b) Con�guration 2

Figure 3: Con�gurations 2 and 3

(a) Con�guration 3 (b) Con�guration 4

Figure 4: Con�gurations 4 and 5

Alshain's �nal con�guration strikes a careful balance between each of the con�guration con-siderations. The vehicle C.G. shifts no more than 0.164 meters during each of the eighteennominal mission scenarios. NASA standards are carefully observed by giving the fore posi-tioned astronaut the pilot controls, which can be monitored by the aft astronaut. Placing theastronauts alongside the propellant tanks limits the z direction C.G. of the vehicle to 2.15meters o� of the ground. The fore astronaut is presented with an excellent �eld of vision thatis limited not by the vehicle, but instead by his or her knees. In the rescue mode, an inca-pacitated astronaut would be loaded upon the rear cargo elevator to accomodate three peoplein total. By removing the landing gear, the stowed area of the vehicle measures 3.2 by 3.8meters. Lastly, with an inert mass measuring 882 kg, Alshain falls well within the reasonablelimits of Altair's cargo capacity of approximately 6000 kg.

The design layout of the Alshain lunar �ying vehicle is the result of a compromise betweeneach of the vehicle's subsystems. No single con�guration is capable of achieving the optimumresult for each consideration, so a careful balance must be struck.

16

Figure 5: Final Con�guration

2.2 Systems Overview

2.2.1 Mission Planning

2.2.2 Avionics

2.2.3 Crew Systems

2.2.4 Loads, Structures, and Mechanics

2.2.5 Power, Propulsion, and Thermal

2.2.6 Systems Integration

2.3 Egress from Altair Lander (Sarah Beal)

A goal of the Constellation program is to create a permanent lunar outpost at the southpole of the Moon near Shackleton crater. NASA's Lunar Surface Systems Project O�ce hasdeveloped a full surface architecture for this lunar outpost to aid with technical work andautomated manual labor. The Alshain Lunar Flying Vehicle is designed to integrate into thecurrent surface architecture of Constellation to improve lunar exploration. Before the AlshainLunar Flying Vehicle can be used for any scienti�c or transport missions, it must be removedfrom the Altair Lander. Alshain is being transported to the moon on an Altair cargo mission.It will arrive on the six meter diameter cargo deck of Altair after the full construction of thelunar outpost. Because the surface architecture will be previously established, crews will beable to take advantage of the tools at base to unload Alshain from the cargo deck.

2.3.1 Resources Onboard Altair

The unloading of Alshain from the cargo deck requires minimal crew involvement and theability to attach the landing gear, which is stowed on Altair detached from the main Alshainvehicle. It also requires the ability to move Alshain away from Altair to the landing pad tominimize the interference from ejecta during take-o�. The necessary hardware componentsinclude the Tri-ATHLETE, the power supply unit, the lunar surface manipulator system, andthe Crew Mobility Chassis.

Tri ATHLETE

The Three-legged All-Terrain Hex-Legged Extraterrestrial Explorer (Tri-ATHLETE) is a unitthat can be operated either autonomously or with minimal crew involvement to unload pay-loads from the cargo deck of the Altair vehicle. Its autonomous operation uses either wirelesslocal area network (WLAN) or S-band communication operating at 150 kbps to provide hori-zon to horizon communication. It has a 13,000 kg payload capacity and can be operated onup to a 30 degree slope, which exceeds the Level 1 Alshain requirement of full operation ona 15 degree slope. It can operate using two di�erent sources of power, either relying on its

17

Figure 6: Altair lander with Tri-ATHLETE and Power Supply Unit stowed on the cargo deck.

(a) Image of a single Tri-ATHLETE unit, consist-ing of three legs.

(b) Image ofa two linkedTri-ATHLETEunits.

Figure 7: Tri-ATHLETE Con�gurations

internal power supply, which is a 6.5 kWh Lithium Ion (Li-ion) battery, or the power supplyunit (PSU). The power supply unit is an external power source also available on all Altaircargo missions. The Tri-ATHLETE comes fully equipped and stowed on each cargo missionas shown in �gure 6 to unload from the payload deck.

A single Tri-ATHELETE unit consists of three legs as shown in �gure 7a. However, it mustbe operated with two connected units to achieve full mobility on the lunar surface as shownin Figure 7b.

The legs of the Tri-ATHLETE have three joints, which can be individually controlled, to mimichip, knee and ankle joints. It has the ability to walk, step, and roll on its wheeled legs. Thelinked Tri-ATHLETE units are one fault tolerant, able to continue if any one of the six legsfail1.

Power Supply Unit

The power supply unit (PSU) can be integrated with two Tri-ATHLETE units to providean extended range of 5 kilometers from the Altair lander. The Tri-ATHLETE units comepre-integrated at its four attachment points with a PSU for unloading from the cargo deck ofAltair. With a PSU connected, the Tri-ATHLETE has its full range of motion of stepping,rolling, and walking on the lunar surface. An integrated PSU and Tri-ATHLETE pair can beseen in Figure 82.

Lunar Surface Manipulation System

The Lunar Surface Manipulator System (LSMS) allows the Alshain Lunar Flying Vehicle tobe removed from the Tri-ATHELETE and placed wherever desired. It has the ability to lift

1Surface Architecture Reference Document (SARD). Ver. 3.4. 2008. p 14.2Surface Architecture Reference Document (SARD). Ver. 3.4. 2008. p 14.

18

Figure 8: Power Supply Unit

Figure 9: Lunar Surface Manipulation System

up to 6 tons. It has a horizontal reach of 7 meters and a kingpost height of 3.75 meters. TheLSMS can be mounted to the cargo bay of the Altair Lander, to the surface of the moon, orto the Crew Mobility Chassis (CMC). Image 9 shows the LSMS mounted to a Crew MobilityChassis3.

Crew Mobility Chassis

The Crew Mobility Chassis, CMC, is intended to move cargo and crew and can tow up to1,400 kg. It has the ability to lift the LSMS and the Alshain vehicle and move it up to 100kilometers. It has the ability to travel on a 30 degree slope, which also meets the Level 1requirement of the Alshain lunar �ying vehicle operating on a 15 degree slope and can moveup to 20 kilometers per hour. It has onboard energy storage provided by a Li-ion battery andcan be driven either autonomously or manually. In case of emergency, it also has the ability toreplenish the oxygen and water in an astronauts PLSS in less than thirty minutes time. Thecrew mobility chassis can be seen in �gure 104.

2.3.2 Mission Plan for Egress from Altair

In order to unload Alshain from Altair, several tasks have to be undertaken. The �rst taskis the physical removal of Alshain from the cargo platform. This will be completed by theTri-ATHLETE unit integrated with a PSU. They must be fully integrated because the egressfrom Altair requires both stepping and rolling motion. The Tri-Athlete's back two legs willstep onto the cargo deck of the vehicle to ensure clearance of Altair's components such as

3Culbert, Chris. �NASA Lunar Surface Systems Project Overview�. NASA, Feb 2009, p 36.4Culbert, Chris. �NASA Lunar Surface Systems Project Overview�. NASA, Feb 2009, p 27.

Figure 10: Crew Mobility Chassis

19

landing gear and RCS thrusters. The front four legs will roll along the lunar surface as theback two roll along the cargo deck. At the end of the deck the back two legs will step ontothe lunar surface and the Tri-ATHLETE will move away from the base.

Once the Tri-ATHLETE has cleared the Lander, the Lunar Surface Manipulator System,mounted to the Crew Mobility Chassis, will connect to Alshain's roll cage and lift it o� of theTri-ATHLETE. It will then move Alshain to the launch pad where the landing gear will bere-attached for �ight. The LSMS and Tri-ATHLETE will then return to the lunar base.

20

3 Mission Planning

3.1 Range Versus Vehicle Mass (Adam Kirk/ Neal Vasilak)

3.1.1 Purpose

The range of the Alshain vehicle is based on the distances to important locations near theplanned lunar base site at the rim of Shackleton Crater on the south pole of the Moon. Theinitial vehicle design was set at a range of 57 km. This distance allowed the vehicle to reachthree scienti�cally important craters around the base site. The craters have been cataloged byNASA with the names Shoemaker, de Gerlache, and Shackleton. Figure 11 shows distancesaway from the base site, giving an idea of what surface features are within the 57 km rangeof the vehicle. It also provides the amount of change in velocity, ΔV, required to reach thatdistance. This one-way ΔV is based on a ballistic trajectory. The basic equations used forthis kind of trajectory can be found in Appendix A.2.

The purpose of the range vs. mass trade study was to get a rough idea of what the masspenalty would be if the range of the vehicle was increased in order to reach further importantsites. Two notable sites would be Malapert, a mountain seen in the top right of Figure 11,and Amundsen, a large crater on the right side of the image. These two surface features areover 100 km away, requiring a signi�cant increase in range to be reachable.

Figure 11: Distances from the base site and the required ΔV

3.1.2 Assumptions and Calculations

The �rst step taken in the trade study was to create a set of increasing propellant tankmasses and �nd the corresponding structural masses required to support a vehicle for thegiven propellant tank sizes. A linear model correlating vehicle propellant tank mass and totalvehicle inert mass was implemented for this calculation.

Once the set of propellant tank and structure masses was compiled, the fuel mass for thevehicle's four spherical tanks was calculated. Using the fuel, structure and tank masses, andan assumed payload mass of 450 kg, the maximum range that this vehicle could travel wasfound using the Tsiolkovsky rocket equation and ballistic hopping equations. Note that themasses of several components such as fuel cells and avionics was held constant, as these scalenegligably to a larger vehicle. One form of the Tsiokovsky rocket equation is given by:

21

Figure 12: Mass vs Max Range Trade Study Plot

=∆initial

final

sp m

mgIV ln

where Isp is the speci�c impulse of the rocket engine, g is earth gravitational acceleration,minitial is the initial mass of the vehicle, and m�nal is the mass post-burn. The total ΔVwas calculated by assuming that the vehicle would make a ballistic hop to its destination andback with no changes in elevation. An additional ΔV of about 100 m/sec was added to eachof the hops to account for fuel loss when searching for a landing location. This amount ofΔV was chosen because it equates to a 3 km glide if horizontal velocity from the ballistichop is retained. It was concluded that this was a conservative amount of ΔV for landing siteselection.

In summary, the major assumptions made for this analysis were:

� Propellant mass scaled linearly with total inert mass

� Rocket engine speci�c impulse of 400 seconds

� Spherical tanks

� 450 kg payload

3.1.3 Results

A plot of the results of running the calculations with the provided assumptions is shown inFigure 12.

As expected, the data showed an increasing trend of inert mass versus range. Because most ofthe vehicle was simply scaled up in size, the mass of the vehicle begins to increase exponentiallywhen straying too far from the initial con�guration range of 57 km. However, the plot is ableto provide a rough, conservative idea of the expected mass increase if the vehicle was scaled toincrease its range to reach further locations such as Amundsen and Malapert. From this plot,it would seem that increasing the range of the vehicle beyond 100 km would be impractical

22

Figure 13: Plot of the relationship between astronauts food speed and stride length.

from a mass standpoint. Thus, it was decided to maintain the range of 57 km instead ofdesigning for an extended range. A MATLAB script was developed to generate the plot seenin the previous section. The full script can be found in Appendix A.2.3.

3.2 Exploration Range (Sarah Beal)

The exploration range from the lunar �ying vehicle is an essential component to the designand implementation of mission plans. This is because the limitations of the exploration rangeare based on the safety of astronauts and thus restrict sites that they will be able to visit.The safety of astronauts is dictated by many di�erent factors such as the occurrence of solarparticle events and illumination area.

3.2.1 Solar Particle Event

In order to ensure safe travel back to base after a strong solar particle event, an analysis wascompleted on the amount of time it takes an astronaut to return to the safety of the lunar basesite. Completing this analysis required knowledge of the anatomy of the astronauts and thelimitation of their motion. The �rst analysis that was completed used Alexander's Formula5:

In this formula, speed is the average foot speed, g is the gravitational constant, stride is thestride length, and hip is the hip height. For a 95th percentile American male, the hip height is0.882 meters and for a 5th percentile American female, it is 0.825 meters6. From this equation,a comparison of speed to stride length could be derived. It is an exponential relationship, anddue to the relatively similar statistics for both men and women, the curves are very similar,as can be seen in �gure 13.

In order to �nd the maximum walking speed on the Moon, the relation for running was used7:

5Math Applications. 2006. DrsCavanaugh. 7 Mar. 2009. http://drscavanaugh.org/digitalcamera/math_applications.htm695 Percentile Person. 7 Mar. 2009 <http://members.shaw.ca/gnat/95.html>.7"Sorby Geology Group." 2003. The University of She�eld. 7 Mar. 2009

<http://www.sorbygeology.group.shef.ac.uk/DINOC01/dinocal1.html>.

23

Figure 14: Exploration range of male and female astronauts as a function of time of �ight.

From this equation it was found that the stride length for the 95th percentile American malewas 2.56 meters, while the 5th percentile female's measured 2.39 meters. The result combinedwith the plot for foot speed verses stride length, it was found that the maximum foot speed forthe male was 1.75 meters per second and the 1.71 meters per second for the female. From thesenumbers, a comparison of the exploration range as a function of time of �ight was created, seenin �gure 14. This demonstrates the distance an astronaut can travel away from Alshain duringscienti�c exploration if they are informed of the solar �are incident thirty minutes before thee�ect of the solar �are reaches the Moon. This allots �fteen minutes for the crew to enter thevehicle at the exploration site and exit the vehicle at the base site. The remaining time, 15minutes, is split into walking time and time of �ight.

As shown from this �gure, the male and female results follow the same trend, however thefemale's walking velocity is the limiting factor. This plot can be used to determine the explo-ration distance an astronaut can achieve at any given range. For example, Shoemaker crater is57 kilometers from the base site, which gives a one way time of �ight of 250 seconds or 4 min-utes and 10 seconds. At this time of �ight, the exploration range for the male is 1140 metersand is 1110 meters for the female. These values represent the shortest exploration distanceof the astronauts because they correlate to the maximum vehicle range of 57 kilometers. Any�ight time shorter than this will allow for farther exploration from the vehicle.

3.2.2 Lighting

The limitations on exploration range for lighting are much more �exible than those for thesolar particle events. The landing lights on the Alshain lunar �ying vehicle have the abilityto illuminate up to a 1.5 kilometer radius. Along with these stationary lights, the crew has aseries of portable lights that can illuminate up to a 10 meter diameter.

24

Figure 15: South Pole elevation map around lunar base site

It is seen that on missions where the time of �ight is very low, around 50 seconds, the limitingfactor on exploration range is lighting rather than solar particle events. However, the closestprojected mission site, Shackleton Crater, has a time of �ight of 145 seconds (10.856 km).Therefore, the time of �ight due to solar �ares will always be the limiting factor for allowableexploration range away from the vehicle.

3.3 Area Surrounding Altair Landing Site

3.3.1 Local Terrain of Lunar Base Site (Adam Kirk)

Lunar Base Site Location

A primary candidate for a future lunar outpost with Project Constellation is on the rim ofShackleton crater, which is located near the south pole with coordinates of about 89.65°S,87°E using the selenographic coordinate system. This location is enticing because it is locatedon a ridge that is continuously exposed to sunlight. In addition, Shackleton Crater has shownstrong evidence for possessing water deposits. Continous sunlight is bene�cial for both humanfactors EVAs, and the ability to provide power through solar cells. Water deposits are soughtafter as a source of hydrogen and oxygen which can be used as propellants.

Terrain Maps

The terrain surrounding the planned lunar base site is fairly rugged and poses some uniquechallenges. Craters are ubiquitous throughout the region with elevation drops as much as 5km. Also, there is a nearby mountain that rises as high as 5 km. Figure 15 shows a colorcoded geographic map depecting areas of various elevation.

The south pole also fails to provide consistant level terrain. The slopes of the terrain typicallyrun between 5° and 15° near the base site. Near the edges of craters it is possible to reachslopes as high as 30°. Rovers such as the T.U.R.T.L.E. are only constructed to handle slopesup to 20°. Figure 16 displays another map of the terrain around the base site with the colorsindicating the magnitude of the slope.

The base's unique position on the South Pole causes the amount of sunlight received to behighly dependent on position. The amount of sunlight in the region is extremely variable anddepends greatly on the relative elevation of the surrounding terrain. This is what allows the

25

Figure 16: South Pole slope map around lunar base site

Figure 17: South Pole % sunlight map around lunar base site

base site to be constantly exposed to sunlight, but also can cause the bottom of craters to beeternally shadowed. Figure 17 provides the percentage of sunlight received near the plannedlunar base sight.

Overall, the south pole of the Moon is a much more challenging region to traverse than themore mare-covered regions encountered by the Apollo astronauts. This is one of the primarymotivations in looking at utilizing �ying vehicles instead of only lunar rovers.

3.3.2 Crater Analysis (Zach Neumann)

Alshain is intended for visiting lunar sites that are normally inaccessible by other means.These include both missions far from the base site and mission locations such as peaks andcraters which involve large changes in elevation. Because of the expected limits on accessiblerange from the base site, focus was placed especially on craters in the area of the lunar southpole (Table 1) as potential mission sites.

Craters are important locations to visit for both scienti�c and logistical reasons. Since cratersare formed by impacts from extralunar bodies, such as meteorites and comets, the crater �oorsoften contain deposits of materials that would not be normally found in the regolith on thesurface of the Moon. Of particular interest are deposits of hydrogen, which have been detectedin the bottoms of certain craters and may indicate the presence of water. Older and largercraters tend to gather larger amounts of volatiles, as they contain smaller craters within them.

26

Table 1: Coordinate Data for Related Sites

Site Name Latitude Longitude

Shackleton 89.9° S 0.0° E Amundsen 84.5° S 82.8° E Cabeus 84.9° S 35.5° W De Gerlache 88.5° S 87.1° W Faustini 87.3° S 77.0° E Malapert 84.9° S 12.9° E Nobile 85.2° S 53.5° E Scott 81.9° S 45.3° E Shoemaker 88.1° S 44.9° E Wiechert 84.5° S 165.0° E �

Additionally, crater �oors that are protected from exposure to direct sunlight, due to craterdepth and location, will contain larger quantities of intact volatile compounds than otherwise.These include water molecules, which would evaporate and be lost if exposed to sunlight. Thecraters near the lunar base site, because of their proximity to the lunar south pole, remain innear to completely perpetual darkness.

The considerations for the selection of potential mission sites for the Alshain vehicle includethe distance from the base site on the rim of Shackleton crater, crater depth, and potentialsigni�cance regarding both scienti�c studies, and potential for in situ propellant mining, whichrelies on the presence of hydrogen or water.

Due to the proximity to the lunar limb, most of the craters near the south pole have notyet been accurately measured in terms of depth. The closest known estimates are givenby calculations based on equations by R.J. Pike8 and the estimates compiled by Dr. JohnWestfall of the Association of Lunar and Planetary Observers (ALPO)9, and the uncertaintiesare unknown.

While planning for the intended range of the Alshain vehicle, the signi�cance of the cratersites, i.e. whether important compounds are detected or suspected within, is weighed againstthe distance from the base site.

Pictures are from the Goldstone Solar System Radar with the exception of the Wiechertphotograph, which was taken by the Clementine spacecraft.

Shackleton Crater

8Pike R.J. (1980) Geometric interpretation of lunar craters. US Geological Survey Professional Paper 1046-C, US GovernmentPrinting O�ce.

9Westfall, John E. 2000. Atlas of the Lunar Terminator. Cambridge Univ. Press.

27

Figure 18: Shackleton Crater

Shackleton crater is the closest potential mission site to the base site, as the site is situatedright on the rim of the crater. Since the rotational axis of the Moon lies in the interior of thecrater, the crater �oor lies in perpetual darkness. This results in a temperature of less than100 kelvin on the �oor, creating a �cold trap� where water vapor and other volatiles withinthe crater are prevented from escaping. In addition, the Lunar Prospector spacecraft detectedhigh concentration of hydrogen in the crater interior. Although tests did not con�rm thepresence of water in the regolith, it is believed that water may still be found in the interior.On the other hand, the crater is relatively young and small compared to other craters nearby,19 km in diameter and 2 km deep, so it may not contain as many volatiles as older and largercraters. Because of the likeliness of the presence of hydrogen and possible presence of water,as well as its adjacency to the base site, makes the crater a certain candidate for mission sitesfor the Alshain vehicle.

Amundsen Crater

Figure 19: Amundsen Crater

Amundsen crater is located near the southern limb of the Moon, facing the Earth. It isapproximately 105 km in diameter, with its rim just reaching around 100 km distance fromthe Shackleton base site. The apparent depth of the crater �oor according to Pike 1980 is 4.06km, while the Westfall estimate for depth is 5.87 km. There is little known about the interior�oor of Amundsen, except that it is mostly �at with a pair of central peaks near the centerof the �oor, and that only the peaks and the southern section of the �oor receiving sunlight.Because of its relative distance from the base site compared to those of the other nearbycraters, as well as the lack of information regarding contents of interest, it was dismissed as apotential mission site for Alshain.

28

Figure 20: Cabeus Crater

Cabeus Crater

Cabeus is located to the northwest of the lunar south pole, west of Malapert crater. Itsdiameter is 98 km, and its midpoint lies approximately 75 km away from the Shackleton basesite. The apparent depth of the crater �oor according to Pike 1980 is 4.03 km, while theWestfall estimate for depth is 5.71 km. As with most other south pole craters, very fewdetails are known about the interior of the crater, but it lies in near-perpetual darkness. Itis not known whether the crater �oor contains any compounds of interest, as none have beendetected at this time, but because of the proximity to the base site, Cabeus is counted amongpotential mission sites, as the darkened �oor may still provide scienti�c opportunities.

De Gerlache

Figure 21: De Gerlache Crater

De Gerlache crater is located just to the west of Shackleton crater, 25 km between the rimand the base site. The crater's diameter is 32.4 km, and the apparent depth of the crater�oor according to Pike 1980 is 2.89 km. The circular crater lies in perpetual darkness, and sothe nature of the interior �oor is unknown. Like Cabeus, de Gerlache is one of the potentialAlshain mission sites, largely because of its convenient placement with respect to the base site,since it is unknown whether the crater �oor contains any useful materials.

Faustini

29

Figure 22: Faustini Crater

Faustini lies directly south of Amundsen, around 75 km from Shackleton crater, midpoint-to-midpoint, with a diameter of 39 km. Like other nearby craters, the �oor of Faustini doesnot receive any light from the sun, and as a result, the crater interior stays permanently at atemperature of less than 100 kelvin. This �cold trap� would mean that any water moleculescontained in the crater, if any, would be trapped within. Although radar observations ofFaustini did not detect any ice in the crater �oor, high concentrations of hydrogen weredetected. Because of its proximity to the Shackleton base site and of the potential for themining of hydrogen (and possibly water) for in situ propellant production, Faustini was chosenas likely mission site for the Alshain vehicle.

Malapert

Figure 23: Malapert Crater

Malapert crater is 69 km in diameter, lies about as far from the lunar south pole as Amundsencrater. As with the other craters, the interior of the crater is unmapped, although the apparentdepth according to Pike 1980 is 3.62 km. The most signi�cant feature of Malapert crater isactually �Malapert Mountain�, a 5-km-high peak on the southwest rim of the crater. Thispeak has been suggested as a site for a communications setup, as it is constantly in sight ofthe Earth, as well as Shackleton crater. Although the site may be revisited at a later date,Malapert crater has been dismissed as a potential mission site for Alshain due to the greatdistance from the base site.

Nobile

Nobile crater is located just to the west of Amundsen. Its diameter is 73 km long, and it liesover 100 km away from the Shackleton base. The Pike 1980 estimate of the crater depth is

30

Figure 24: Nobile Crater

3.69 km, while the Westfall estimate is 3.74 km. Sunlight enters the interior of the craterobliquely when it enters at all, and the interior of the crater is irregular, with smaller craterswithin. Since there is no known indication of minable materials, and because of the distancefrom the base site, Nobile is not considered as a potential mission site for Alshain.

Scott

Figure 25: Scott Crater

Scott lies north of Nobile, just less than 200 km away from Shackleton. The Pike 1980 estimateof the crater depth is 4.09 km, while the Westfall estimate is 5.57 km. Because of its proximityto the lunar limb and the south pole, the northern end of the crater lies in near-perpetualdarkness. Because of the prohibitive distance from the south pole and the base site, Scott wasnot considered as a potential mission site for Alshain, especially since there have not been anymaterials of particular interest, particularly hydrogen, detected in the crater �oor.

Shoemaker

Shoemaker lies just to the west of Faustini and the south of Malapert, about 50 km fromShackleton, crater-to-crater. The apparent crater depth is 3.31 km according to Pike 1980.Like in the case of Faustini, the lack of illumination on the crater �oor creates a �cold trap,�an enclosure with a temperature below 100 kelvin, which traps any water molecules that arepresent in the crater interior. Also similar to Faustini, high concentrations of hydrogen weredetected by the Lunar Prospector. Although tests have failed to con�rm the presence of water

31

Figure 26: Shoemaker Crater

in the �oor of Shoemaker, the hydrogen can still likely be extracted for use as propellants.Because of the presence of the hydrogen deposits in the crater �oor, as well as the positionrelative to the Shackleton base, Shoemaker is a likely prospect as a mission site for the Alshainvehicle.

Wiechert

Figure 27: Wiechert Crater

Weichert crater is located on the far side of the Moon, less than 170 km away from thesouth pole. Approximately 41 km in diameter, the crater is highly eroded and illuminatedby oblique sunlight, resulting in parts of the interior lying in deep darkness. The apparentdepth is approximately 3.10 km according to Pike 1980. There is very little known of interestabout the contents of the interior �oor, and it is located prohibitively far from the Shackletonbase site, farther than the other large craters under consideration. Therefore this crater wasdismissed as a potential mission site for Alshain.

Conclusions

In conclusion, the craters that were selected as possible mission sites for Alshain are Shack-leton, Cabeus, de Gerlache, Faustini and Shoemaker. These were the sites that were withinthe chosen range of 57 km, and some of them are suspected to contain hydrogen and othermaterials of interest.

3.4 Payload Bay Sizing (Fazle Siddique)

The Alshain payload bay has a contingency requirement to support a downed astronaut inemergency situations. For the payload bay to be compatible with an upright-seated 95th-percentile male astronaut, the platform must support payload weights of 170 kilograms and apayload platform area of 0.86x1.0m with a clearance of 1.49m (volume of 1.28m3).

32

Using these payload bay parameters as a baseline, it was essential to analyze whether theseparameters would allow the Alshain payload bay to be adequately compatible with commonscienti�c payloads. Thus, a database of 104 payloads was constructed from NASA documentsconsisting of past Apollo science payloads and future science payloads slated for use on theMoon and Mars (see Appendix A.3). The study analyzed how many of the individual payloadswould be compatible with the payload bay for a given mass or volume limit. The aim was toconclude whether mass and volume limits dictated by diminishing returns on the percentage ofcompatible payloads were within the range of the parameters dictated by a downed astronaut,or if augmentation of the payload bay was necessary. All calculations and plots were completedin MATLAB (see Appendix A.2.7).

Figure 28 is the plot of the percentage of payloads compatible with a payload bay of increasingmass capacity. Point of diminishing returns would dictate a 40 kg payload bay capability. Notethis is only for an individual instrument, and thus having a payload bay capable of supporting170 kg will allow for several scienti�c instruments to be carried, with additional room forregolith. Also note that in the event of an emergency the scienti�c payload will be o�oadedin order to make accomadations for an injured crew member who is unable to reach his/herseat.

Figure 28: Payload Bay Mass Plot

The same can be seen with the volume of payload bay versus percentage of compatible payloads(Figure 29). Point of diminishing returns would dictate a 0.28 cubic meter payload bay volume,well within the range of the requirements of supporting a downed astronaut. This is alsomaking the assumption of a single scienti�c instrument, but a 0.28 cubic meter volume ascompared to the designed 1.28 cubic meter allows for several instruments to be mounted andcarried on each �ight.

33

Figure 29: Payload Bay Volume Plot

3.5 Locking Mechanism for Payload Bay (Sarah Beal)

In order to successfully and safely secure each item into the payload bay, an adaptive methodhad to be designed which allowed for payloads of di�erent shapes, sizes and weights to beeasily stowed and removed by an astronaut while on EVA. In the case of this payload bay, the�oor will be made out of aluminum isogrid. It contains a series of equilateral triangles intowhich mushroom locking mechanisms will �t to secure the payloads. This is a useful techniquebecause the payload will not only be placed in a predetermined location to balance the centerof gravity of the payload bay, but their �xed location will also stop any payload movementduring �ight that would cause a shift in center of gravity.

The isogrid material has a high strength to weight ratio, which makes it ideal for use inthe payload bay where a minimal mass is required to support the weight of the payloads.The maximum payload that will be applied to this payload area is that of an incapacitatedastronaut. In this case, a standard PLSS locking attachment set will exist on the elevatorrails. By locking their PLSS into thie, attachment and using foot constraints, the astronautwill remain within a 1 x 0.86 meter �oor area and a height of 1.44 m for a 95th percentileAmerican male. The male will weigh 170 kilograms, which means that the isogrid must beable to stand at least this much weight.

3.6 Refueling Alshain (Andrew Wilson)

As with many types of vehicles the Alshain Lunar Flying Vehicle will need to be refueledwith each use. Part of mission planning requirements was that Alshain must be designed forsimple interfacing for consumables re-supply, maintenance, and servicing. The assumptionsmade to design for the refueling of Alshain were that there would be in situ propellants on theMoon as well as a propellant processing station provided by the Constellation Lunar SurfaceArchitecture. This station would have both the capability of processing residual propellantsfrom the Altair Lunar Lander as well as the in situ propellants on or within the lunar surfaceand all the necessary equipment to transport cryogenic propellants to and re-supply Alshain.

Proposed here is that an Alshain landing site be located as close as possible to a propellantprocessing station to provide quick and non-hazardous access for optimal refueling and mainte-nance EVAs. The use of a landing pad or surface is suggested to protect the propellant station

34

from hazardous ejecta upon use of Alshain nearby. Considering Alsahin's overall dimensions,a 10x10 meter area that is well lit and near the lunar base would provide a suitable landingsite well within Alshain's accuracy capabilities.

One material considered for the landing surface was ceramic that could be machined into tilesto be transported with the Altair and lain out by astronauts while unloading Alshain. Commonceramic tiles such as titanium or silicon carbide (TiC or SiC) have melting temperaturesbetween 2000 � 3500 Kelvin which is well beyond the 700 K temperature experienced at themouth of the engine. Unfortunately they would not be mass e�cient to transport to theMoon, weighing in at approximately 2000 � 3000 kilograms for a 100 square meter area, aswell as being prone to crack substantially upon landing, e�ectively producing ejecta insteadof preventing it.

Therefore a simple �re blanket commonly used by �re�ghters was deemed a suitable landingsurface because of its ability to remain intact under temperatures comparable to that ofthe ceramics and be easily secured to the lunar surface beneath Alshain upon unloadingfrom Altair. Speci�cally, the Insul�ex Pyroblanket 17oz made of silicone rubber would weighapproximately 60 kilograms for a 100 square meter area.

3.7 Dust Maintenance (Adam Kirk)

During the Apollo missions, the accumulation of lunar dust on astronauts, equipment, andvehicles posed a problem. It was found that the lunar dust tends to cling to astronauts andequipment and is very abrasive. For the short periods that the Apollo missions were conducted,this was more of an annoyance than a major issue. However, with Project Constellation's planto have a long-term base on the Moon, the issue will have to be addressed. Because lunar dusthas an electrostatic charge, it is likely that an electrostatic device will be developed to repelthese dust grains and clean equipment. For instance, one concept that has been looked intois an �electric curtain� composed of parallel electrodes that could sweep such particles o� itssurface10.

For Alshain, it would be necessary to periodically remove dust from important componentsafter missions with an electrostatic device. Key areas would include control panels, the mainengine nozzle, reaction control system nozzles, PLSS connectors, and the payload elevator.Testing would be required to determine how often maintenance would be necessary.

10http://science.nasa.gov/headlines/y2006/19apr_dustbuster.htm

35

4 Propulsion System

4.1 System Overview (Nitin Sydney)

The propulsion system of the Alshain is designed to e�ciently meet all the level one require-ments. In addition, several other requirements were created by the design team to completelyensure astronaut safety during a nominal mission and 24 hour contingency operation. TheAlshain propulsion system is shown below.

Figure 30: Propulsion System Overview

There are two LOX and two LH2 tanks holding a total of 940 kg of fuel. This is used forpropellant for the main and RCS engines, and for the PEM fuel cells (not shown in thediagram). The LOX can also be used in contingency operation for oxygen for the astronauts.The fuel is fed to various systems using a pressure feed system pressurized at 8.4 MPa. Thereare four pressure tanks cross-fed to a two-fault tolerant pressure regulator system which scalesthe pressure to 2.4 MPa. In the event of a pressure loss from one of the pressure tanks, thethree remaining tanks can handle all the pressurization requirements.

There is one main engine and 20 RCS thrusters onboard the Alshain. The main engine operatesat a combustion pressure of 2 MPa and a thrust of 40 kN. The RCS engines also operate at acombustion pressure of 2 MPa, but have thrusts varying from 300 to 1150 kN. The propulsionsystem is designed so that if the main engine fails to reignite, the RCS system will take overand land the crew safely.

The entire system is regulated by two-fault tolerant pressure regulators and propulsion valves.In addition there are mechanical �ll and drain valves for refueling purposes. In the event of afailure in the pressure regulator system, there are mechanical pressure relief valves attachedto each tank to prevent catastrophic failure modes.

4.2 Main Engine System (Matt Kosmer)

Design of the Main Engine System (MES) on the Alshain was performed using a combinationof thermodynamic relations for an ideal rocket and historical examples of LOX/LH2 engines.Because existing LOX/LH2 engines are much too large for use with a relatively small lunar

36

�ying vehicle, we had to estimate the characteristics of an engine built speci�cally for thisapplication. The resulting Main Engine System consists of a single 40 kN engine, 100 cm tallby 80 cm in diameter, with an Isp of 400 seconds, expansion ratio of 45:1, and a chamberpressure of 2.0 MPa.

4.2.1 Thrust Requirements

There were three main concerns when determining the total thrust capability of the MainEngine System: safety of crew, minimizing waste of fuel due to gravity drag, and minimizingengine mass. The following analysis shows how these three factors resulted in the decision tomake an engine capable of 40 kN of thrust.

Crew safety was a limiting factor in the engine design that set an upper limit on how muchthrust the engine could produce. In order to prevent injury to the crew of the Alshain, amaximum acceleration of 2 g's (19.6 m/s2) was set. Estimates of the dry-mass and fueled-mass of the vehicle were then used to determine an approximate range of acceptable values ofthrust for the engine. Fully fueled at approximately 2500 kg, the Alshain would achieve 2 g'swith approximately 49 kN of thrust. Similarly, with a dry mass of approximately 1500 kg, 29kN of thrust would achieve 2 g's. Therefore, the Main Engine System must be able to throttleand have a nominal thrust value between 29 and 49 kN. Anything higher than 49 kN wouldbe unsafe for the crew of the Alshain.

The next factor to be analyzed was the e�ect that increasing engine thrust had on reducinggravity drag. Gravity drag is the term given to the reduction in ΔV capability of an enginedue to the presence of a gravity �eld. This reduction in ΔV is directly proportional to boththe strength of the gravity �eld and the burn time of the thrusting maneuver as shown in thefollowing equation, a modi�ed version of the rocket equation:

∆V = −Ve ln

m f

mi

− g∆t

where Ve is the exit velocity of the engine, mf and mi are the �nal and initial masses ofthe vehicle respectively, g is the acceleration due to gravity, and Δt is the burn time. Thecomponent gΔt is the loss inΔV due to gravity drag. The goal of this analysis is to determine,with a known �nal mass of the vehicle and target ΔV, how much fuel is required to achievethe necessary ΔV.

Minimizing burn time is a clear way to reduceΔV losses from gravity drag but the relationshipwith thrust is highly non-linear. Burn time is a function of the amount of fuel and the mass�ow rate of that fuel through the engine. The mass �ow rate is simply the thrust dividedby the exit velocity of the engine. Combining these relations with the previous equation, theequation for ΔV as a function of vehicle mass and thrust is as follows:

∆V = −Ve ln

m f

mi

− gVe

Tmi − m f( )

wherein the value of initial mass (�nal mass plus fuel) must be solved for iteratively for eachvalue of thrust, T. This relationship shows that the mass of fuel required to achieve a certainvalue of ΔV decreases with increasing thrust but in order to achieve mass e�ciency for thevehicle as a whole, the mass of the engine must be accounted for.

37

The mass of the engine was estimated by collecting data on thrust versus mass for historicalLOX/LH2 engines such as the RL-10 series, HM-10 series, RD-56, and HM-7 among others11.From this data, a mass estimation of 2.4 kg for every kN of thrust was used in the designof Alshain's engine. When this linear mass estimation relationship is combined with thedetermination of fuel mass from the gravity drag equation, a clear point of minimum systemmass is found. This process was repeated iteratively as the engine design (expansion ratio, Isp)and the vehicle's dry mass changed during the course of the design process, which resulted ina �nal thrust value of 40 kN as is shown in the �gure below.

System Mass vs. Thrust

940

960

980

1000

1020

1040

1060

1080

1100

1120

0 10 20 30 40 50 60 70

Thrust (kN)

Fu

el +

En

gin

e M

ass

(kg

)

Figure 31: System Mass vs Thrust

4.2.2 Number of Engines

Our rationale for choosing a single engine for the Main Engine System of the Alshain stemsfrom an analysis of the reliability of multiple engine systems versus single engine systems andassumes that any single engine can be developed to meet minimum reliability standards setforth by NASA. The key conclusion that can be drawn from this analysis is that every engineintroduced into the system increases the likelihood of an engine failure event and, as such, asingle engine system will be more reliable than a multiple engine system.

For a multiple engine system, the reliability of the system as a whole, which is to say theprobability that there are no engine failures, is governed by the equation:

P = RN

in which R is the reliability of a single engine and N is the number of engines in the system.It can be quickly gleaned from this equation that as the number of engines increase, theprobability that the entire system of engines is working will decrease. However, if the enginesremaining after a failure are capable of producing the same amount of thrust as a single enginesystem, then we can consider the failure of one engine to still constitute a successful �ight.The probability that either all engines work or only one fails is given by the following equation:

11http://www.astronautix.com/props/loxlh2.htm

38

P = RN + NRN (1− R)

This equation results in reliabilities at least as high as the single engine reliability and thereforea multiple engine system can be considered as reliable as a single engine system as long as theremaining engines after a single failure are capable of producing the nominal thrust value (40kN).

The major downside of a multiple engine system comes from the added size and mass of theredundant engine. The thrust required by each engine in a multiple engine system as a resultof this redundancy is:

T = 40

N −1

When this is combined with the mass estimating relationship developed previously, the totalmass of a multiple engine system is governed by:

M = 2.4

40N

N −1

Because the mass of a single engine system is equal to 2.4 times 40 and N/(N-1) can never beless than 1, a multiple engine system can never weigh less than a single engine system underthese constraints. As a result of the excess mass caused by a multiple engine system, theAlshain has been designed as a single engine vehicle.

4.2.3 Selection of Performance Characteristics

The next step in the design of the Main Engine System was to determine the values of keycharacteristics that would lead to the most desirable performance of the engine. After muchanalysis using thermodynamic relations, the performance and size of the engine was found todepend on two controllable factors: expansion ratio and chamber pressure. Key assumptionsabout the conditions in the combustion chamber were made based on examples of existingLOX/LH2 engines and included a combustion temperature of 3600 K and an oxidizer to fuelmixture ratio of 6:112. After these assumptions were made, thrust, chamber pressure, and Isp(a direct function of expansion ratio, as will be shown) were the only design choices that had tobe made and the rest of the information about every point along the engine from combustionchamber to throat to nozzle exit could be derived from thermodynamic relations.

The �rst calculation was the determination of the chemical properties of the exhaust gas basedon the mixture ratio of 6:1. Balancing the equation of the combustion of LH2 with LOX underthis mixture ratio yields an exhaust gas consisting of water and gaseous hydrogen in a molarratio of 3:1. The molar mass of the exhaust gas is thus 14 kg/mol. The ratio of speci�c heats,γ, for the exhaust gas is calculated using the degrees of freedom method:

γ = f + 2

f

12http://www.astronautix.com/props/loxlh2.htm

39

in which f is the number of degrees of freedom of the gas. Gaseous water, which has 9 degreesof freedom, has a γ of 1.22 and gaseous hydrogen, with 5 degrees of freedom has a γ of 1.4.Averaged together at a molar ratio of 3:1, the exhaust gas of the rocket engine has a γ of1.26. From this, the speci�c heat, Cp, of the exhaust is calculated to be 2878 J/kg/K via theequation:

Cp = Rγ

M(γ −1)

where R is the ideal gas constant and M is the molar mass of the exhaust.

The next step was to use thrust and Isp to calculate the exhaust velocity and mass �owthrough the engine. Thrust was previously calculated to be 40 kN and an array of values ofIsp is used to determine how Isp a�ects the size and dynamics of the engine. The e�ectiveexhaust velocity of the engine, by de�nition, is simply Isp multiplied by the acceleration due toEarth's gravity. Mass �ow is then determined by dividing thrust by e�ective exhaust velocity.Along with our choices of chamber pressure (arranged in an array similar to that for Isp), thethermodynamic conditions of the combustion chamber are now fully determined.

Next the conditions of the nozzle throat must be determined. First, the pressure at the throatis determined by the critical pressure at which the �ow becomes sonic and is determined bythe following equation:

Pt = P0

2

γ +1

γγ −1

in which P0 is the chamber pressure. From this information, the temperature of the gas atthe throat is determined by:

Tt = T0

Pt

P0

γ −1

γ

where T0 is the combustion temperature. The area of the throat is then calculated usingthese values of pressure and temperature as well as the mass �ow, mdot, via the followingrelationship:

t

t

tt T

M

R

MP

RTmA γ&=

which completes knowledge of the state of the engine throat.

The nozzle exit conditions are determined by �rst calculating the exit temperature based onthe previously determined exhaust velocity:

Te = T0 − Ve

2

2Cp

40

where Cp is the speci�c heat of the exhaust. From the exit temperature, exit Mach numbercan be derived using the following relationship:

Me = 2

γ −1

T0

Te

−1

An important note about exit Mach number and exit temperature is that they are independentof chamber pressure and vary only with Isp and thrust. As such, Isp is intimately tied to theexpansion ratio, the ratio of exit area to throat area, ER, which is calculated using the followingrelationship:

ER = M t

Me

1+ γ −1

2Me

2

1+ γ −12

M t2

γ +1

2 γ −1( )

in which Mt is the Mach number at the throat and by de�nition equals 1. The area of thenozzle exit can be found simply by multiplying the area of the throat by the expansion ratio.

Because the expansion ratio depends only on thrust and Isp, this means that physically chang-ing the ratio of the nozzle exit area to nozzle throat area changes the Isp of the engine. Thisleads to making a design choice about the appropriate expansion ratio. Based on the �g-ure below, we chose an expansion ratio of 45:1, which corresponds to an Isp of 400 seconds.We selected an expansion ratio at the lower end of the knee in the curve because increasingexpansion ratios yield quickly increasing nozzle diameters and size becomes a restriction.

Isp vs. Expansion Ratio

350

360

370

380

390

400

410

420

430

440

0 100 200 300 400 500

Expansion Ratio

Isp

(sec

on

ds)

Figure 32: Expansion Ratio vs ISP

Similarly, size played an important role in determining the chamber pressure of the engine.Chamber pressure has a linear e�ect on the mass of the propellant and pressurant tank system,which would lead us to choose the lowest chamber pressure possible. However, at low chamberpressures the engine becomes exceedingly large as can be seen in the �gure below and as such,a moderate chamber pressure of 2.0 MPa was selected.

41

Nozzle Diameter vs. Chamber Pressure

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

0 2 4 6 8 10

Chamber Pressure (MPa)

No

zzle

Dia

met

er (m

)

Figure 33: Engine Diameter vs Engine Pressure

The length of the nozzle was determined by modeling it as a bell nozzle, which featurescomparable expansion e�ciency to a 15-degree half-angle conical nozzle but with only 75% ofthe length, using the following equation:

L = 0.75

re − rt

tan 15°( )

in which re and rt are the exit and throat radii respectively. This is the �nal calculation indetermining the characteristics of our engine which are as follows: Thrust � 40 kN, Diameter� 80 cm, Length � 100 cm, Expansion Ratio � 45:1, Isp � 400 seconds, Chamber Pressure �2.0 MPa, Mass � 96 kg.

4.3 Propulsion Analysis (Nitin Sydney)

4.3.1 Analysis Requirments and Assumptions

The goal of the propulsion analysis is to �nd the minimum amount of propellant that Alshainneeds to complete its max round-trip mission distance of 114 kilometers. One of the require-ments for the vehicle is that it must be able to use in-situ propellants to operate. For themoon, this means that the Alshain must use liquid oxygen (LOX) as the oxidizer and liquidhydrogen (LH2) as the fuel.

The Alshain has speci�c ΔV requirements for the max mission distance. For each hop willtake 700 m/s and there is an additional glide ΔV of 200 m/s for picking the landing site. Thisgives a total ΔV of 1600 m/s. In addition, there must be extra fuel added for the PEM fuelcells, the RCS system, and to account for gravity drag (a non-impulsive burn penalty).

From the mass budget, the approximate inert mass of the vehicle is 1100 kg with a 30% margin.The propulsion analysis is done with this 30% margin added on, so all the propellant valuesare conservative estimates of what the vehicle will actually need. In addition to this inertmass, the Alshian must also be capable of carrying two astronauts plus additional payload.This additional payload must be equal to at least the weight of an additional astronaut. Thisgives a total payload requirement of 500 kg.

42

4.3.2 Propellant Requirements

To �nd the propellant requirements of the vehicle, the max-distance mission must be split intotwo separate parts: This initial ballistic hop and hover and the return ballistic hop and hover.For each of these legs, the rocket equation must be applied to �nd the fuel required from eachleg. The �nal mass for the second leg is the inert mass plus the payload mass, which is 1600kg. Since the ΔV requirement for the second leg is 800 m/s and the engine speci�c impulseis 400 s, the rocket equation dictates:

Where MR is the mass ratio, Isp is the speci�c impulse of the engine, and g is the gravitationalconstant. Using this the initial mass of the second leg can be found by:

Where Mf,2 is the �nal mass of the second leg and Mi,2 is the initial mass. This initial massis also the �nal mass of the �rst leg, . This process can then be repeated for the �rst leg:

This gives a �nal vehicle mass of 2300 kg. Subtracting the inert mass and payload of the vehiclefrom this �nal mass gives the total propellant mass to be 700 kg. But this doesn't account forthe fuel requirements of the RCS system, the fuel cells, and the gravity-drag penalty.

For the RCS, the mass �ow rate for the larger engines, which will be �ring during �ight, isapproximated to be 1.2 kg/s. The total burn time for the main engine is 80 seconds, whichleads to a total additional RCS fuel mass of 100 kg. The fuel cells only require 3 kg for thenominal 8 hour mission, but they will have a mixing ratio of 10:1 rather than the 6:1 thatthe main engine requires. The gravity drag penalty is solved for iteratively to give a �nalpropellant mass of 940 kg. This information is summarized in table 2.

Table 2: Gravity Drag Penalty

Fuel Uses Mass [kg]Flight 700RCS 100Fuel Cells 3Gravity Drag 137Total 940

43

This gives a �nal vehicle mass of 2600 kg. The vehicle masses for this max mission distanceare summarized in the following table.

Table 3: Max Mission Masses

Mass [kg]Initial Mass 2600Final Mass 1600Second Leg Initial Mass 2120Total Propellant 940First Leg Propellant 520Second Leg Propellant 420Inert Mass 1100Payload Mass 500

4.4 Tank System (Nitin Sydney)

4.4.1 Tank Requirements

For the tank system, there are several requirements. First are the safety requirements. Sincethe Alshain is a small vehicle, the propellant feed system is pressure tanks. This means thatall the tanks will be pressurized. Therefore, for astronaut safety, the tanks must have highsafety factors. For low pressure tanks, such as the propellant tanks, a safety factor of two isused, and for high pressure tanks, such as the pressurant tanks, a safety factor of three is usedin accordance with NASA standards.

The tanks are designed to meet several requirements. First is that the tanks use LOX andLH2 as propellants, and have a mixing ratio of 6:1. This is used to obtain the volume of thepropellant tanks. Using that, the volume of the pressure tanks can be chosen to minimize thetotal mass.

The second requirement is to minimize the mass of the tank system. This includes the massof the pressure tanks, the pressurant (which is essentially inert mass for the vehicle) andthe propellant tanks (not including the propellant as that is a set constant). Several variablescontribute to this �nal mass. This includes the number, the shape, the reliability requirements,and the overall center of gravity shift of the tanks. The number and reliability of the tanks areclosely tied. For the shape of the tanks, two options were considered, spherical and cylindrical.

The third requirement is to minimize the vehicle volume. This requirement is secondary tominimizing the mass, but it is still an important consideration. Alshain must �t on the deckof the Altair lander, which to the current knowledge of the design team is 6 by 8 meters.Although this is the size of the deck, the vehicle must be signi�cantly smaller than this so thatother payloads can be send in addition to the Alshain.

4.4.2 Propellant Tanks

Material Choices

44

The material for the tanks is chosen to minimize the overall mass. This is accomplished bychoosing the material with a high strength to weight ratio. However, since the propellantsare at cryogenic temperatures, not all materials are available. Below is a table summarizingseveral tank materials.

Table 4: Propellant Tank Material Summary

Material Density [kg/m^3] Strength to Weight RatioSteel 690 7800 88400Aluminum 400 2700 148000Titanium 830 4510 184000Carbon Composites 1000 1900 526000

Yield Stress [Mpa]

From these materials, aluminum is the most commonly used material for cryogenic tanks andhas a technology readiness level of nine. But the strength to weight ratio is the second lowest.Titanium and carbon composites give the highest strength to weight ratios, but have thelowest TRLs for cryogenic tanks. Titanium has been proven in lab settings for LOX, but notfor LH213. However, there have been recent tests with composite tests that have shown thatthis material has the capability of being space rated within the next ten years14.

To access the mass savings for choosing titanium or composites, a basic study comparing themass of one spherical tank was done. The following table shows the mass of one spherical tankfor each material at two MPa.

Table 5: One Spherical Tank Mass

Material Tank Mass [kg]Steel 21.3Aluminum 12.7Titanium 10.2Carbon Composites 3.6

This table shows that carbon composites are only 25% the mass of titanium, the next strongestmaterial. For this reason, carbon composites have been chosen to use as the material for boththe pressurant and propellant tanks.

Spherical vs. Cylindrical

The primary factor in choosing spherical vs cylindrical tanks is minimizing the mass of thesystem. First the volume of each tank must be found. This is found by splitting the propellantinto oxidizer and fuel with a six to one ratio and dividing the masses by the respective densities.Then, assuming thin walled tanks, the mass of each tank can be found using one of the followingtwo relations:

13irfu.cea.fr/Phocea/�le.php?class=std&&�le=Doc/Publications/Archives/stcm-01-05.pdf14http://www.netcomposites.co.uk/news.asp?2390

45

The �rst equation is the mass of a spherical tank and the second is the mass of a cylindricaltank. P is the pressure in the tank, V is the volume, ρ is the tank material density, is theyield stress, R is the radius and L is the length. The total volumes and propellant masses arelisted below.

Table 6: Spherical Tank Volumes and Masses

The radius and length of the cylindrical tanks were chosen to minimize the tank mass. Shownbelow is a plot of the total mass of the propellant tanks, with insulation, versus number oftanks for each shape.

For making this graph, the pressure is each of the tanks was 2.4 MPa, which is the combustionpressure of the engines plus an extra 20% pressure to account for losses across the controlvalves and the injector plates. Also, with each tank there is an additional 3 kg added toaccount for structural connections to the vehicle. Note that the vehicle size would have toincrease with the number of tanks. This extra structural mass is not contained in this graph.

1 2 3 4

0

20

40

60

80

100

120

Tank Mass Study

Spherical

Cylindrical

Number of Tanks

To

tal P

rop

ella

nt T

an

k M

as

s [k

g]

Figure 34: Propellant Tank Characteristics

From this perspective, it is plainly seen that spherical tanks are better from a mass standpoint.For this reason, spherical tanks are used on the Alshain. It is seen here that the tank massvaries linearly with number of tanks for up to four tanks of each type of propellant. This

46

means that the less propellant tanks will have a lower mass. However, in order to balancethe C.G. of the vehicle, at least two tanks of each are needed. Therefore, two tanks of eachpropellant are used on the Alshain. The characteristics of the tanks are shown in the followingtable.

Table 7: Tank Mass Study Chart

4.4.3 Pressurant Tanks

There are three primary driving factors in designing the pressure tank system. First is themass of the system. The pressure tanks are the heaviest tanks, and therefore their mass mustbe reduced as much as possible. The second driving factor is reliability. The tanks themselvesare assumed to be 100% reliable. This is reasonable since there is a safety factor of threeattached to each tank, but it cannot be assumed that the high pressure lines are that reliable.Therefore, in order to reduce the possibility that there could be a loss of pressure, there isone extra pressure tank aboard the Alshain. Third, is the secondary volume constraint. Thenumber and shape of the tanks must be chosen to reduce the area footprint of the vehicle.

Mass Minimization

For spherical pressurant tanks, there is a closed-form solution for the minimum mass of apressure tank that must �ll a certain volume. To show this, the system is simpli�ed to thefollowing model:

Propellant Tank with minimal pressurant

Tank System at Beginning of Flight

Propellant Tank after mission

Tank System After Flight

Figure 35: Mass Minimzation Model

47

In both of the cases, the pressure in the propellant tank must stay at the combustion pressureof the engine. Assuming an isentropic expansion from the pressure tank to the propellant

tank, the following relation can be used: PV γ=const . . This can be written for stages shownin the above �gure. Since this must be a constant, both equations must be equal to each other,yielding the following relationship:

γpr

ppr +V

V=

P

P 10

Where Ppr is the initial pressurant pressure, Vp is the propellant tank volume, Vpr is thevolume of the pressurant tank and gamma is the speci�c heat ratio of the pressurant. Solvingfor the initial pressurant pressure in the previous equation and substituting into the equationfor the mass of a thin-walled sphere yields:

σρV+V

VP=M pr

γpr

p 12

30

To minimize this mass in terms of volume,

dV

dM

was found and set equal to zero. This yields:

1−γ=

V

V

p

pr

. However, this means that that mass of a tank is independent of number of pres-surant tanks. This is because the pressurant tank volume is simply split into the number ofpressure tanks attached to a single propellant tank. Since M varies linearly with volume, ifthe volume is cut in half, then the mass is cut in half and the total pressure tank mass staysconstant. Note that this does not include any structure mass or the mass of a spare tank.From this, the mass of the system can be shown versus the number of pressure tanks.

Each of these masses includes a spare tank. This is why the mass decreases as the number oftanks increases. The point of diminishing returns is around four tanks, so that is the numberof pressure tanks chosen for the Alshain

However, due to C.G. and sight-line purposes, the tanks were chosen to be cylindrical. Thenumbers shown on the graph above are for such cylindrical tanks. The relationship derivedabove cannot be done analytically for spherical tanks. This is because the minimization ofthe tank mass for cylindrical tanks is an NP complete problem, meaning there is an in�niteamount of solutions that minimize the tank mass. But, as can be seen in the above �gure, thedecreasing mass trend is still valid for cylindrically shaped tanks.

Pressurant Choice

There were two options that were considered for the choice of pressurant: gaseous nitrogenand gaseous helium. Both are inert gases and are commonly used pressurants. Note that theseare not in-situ materials, but since the propellant feed system is closed, the pressurant can bemostly conserved. Minor pressure leaks will eventually cause a loss in pressurization, so there

48

1 2 3 4 5 6 7

0

50

100

150

200

250

300

Total Tank Mass

Total Tank Mass

Number of Tanks

Tota

l Tan

k M

ass

[kg

]

Figure 36: Total Tank Mass

will have to be additional pressurant brought on the Altair lander. A trade study was done to�nd the mass e�ect of either pressurant on the system. This can be done assuming that theideal gas law holds for the regime of interest, which it does. The results are presented in thetable below.

Table 8: Pressurant Choice

Helium NitrogenSpecific Heat Ratio 1.4 1.66Molar Mass 4 32Pressurant Mass [kg] 42 313

Clearly helium is a better option for pressurant. However, it must be noted that while heliumis the mass e�ective option, it has a higher likelihood of leaking. For this reason, the reliabilityof the system would go down, but this is a worthwhile trade o� since at least one pressuretank is already added to the system to increase reliability.

4.4.4 Tank System Summary

To summarize the tank system, the following table includes all the masses, dimensions andpressures for the tanks.

Table 9: Tank Summary

49

This gives a total inert mass of 213 kg for the tank system.

Table 10: Tank Summary

4.5 Valves, Pressure Lines and Pressure Regulators (Nitin Sydney)

4.5.1 Valves and Pressure Regulators

The driving design constraint for the valve and pressure regulator system is two-fault tolerance.If any two valves fail, the system must still be fully functional. Because of this, the followingvalve and pressure regulator system is implemented:

Figure 37: Valve and PR System

With this system, any two faults are accounted for. The system starts by using the right mostvalve and moves left. This will work whether a valve or pressure regulator sticks closed oropen.

The mass and power for each valve and pressure regulator are calculated using estimatingrelations. The power estimate includes heaters for each valve so that the mechanisms do notfreeze due to the cryogenic fuel running through them. The results of the estimations areshown in the table below.

50

Table 11: Valve and Pressure Regulator Estimations

Pressure Regulators ValvesNumber 3 12Power Each [W] 5 5Mass Each [kg] 2 1Total Power [W] 15 60Total Mass [kg] 6 12

In addition to these electric pressure regulators, each propellant tank is out�tted with amechanical pressure regulator that is diagrammed below.

Figure 38: Mechanical Pressure Regulator

If the pressure in any of the propellant tanks passes a critical value, the linear spring willcompress and vent the gas into the lunar vacuum. While this is a loss of pressurant, it avoidsthe failure of tanks. This is simply a measure added on top of the two-fault tolerance of thepressure regulators to ensure astronaut safety.

In addition to the propulsion valves listed above, there will be additional valves attachedto every tank for �lling and draining. The masses of these valves are included in the massestimates for each tank given before.

4.5.2 Pressure and Propellant Lines

Since the pressure and propellant tanks are assumed to be perfectly reliable, the high pressurelines are allowed to fail. As mentioned earlier, in order to increase mission reliability, an extrapressure tank is added to the system. However, extra propellant tanks have not been addedas this would, at the very least, double the amount of propellant carried aboard the Alshain.So, to minimize the risk of fuel loss, all the tanks and propellant tanks are cross-fed as shownin the System Overview section earlier.

51

A simple mass estimation was used to �nd the mass of the propellant lines. The results of theestimation are tabulated in table 12.

Table 12: Pressure and Propellant Lines

High Pressure Line Low Pressure LineEstimated Length [m] 13 15Estimated Mass per Meter [kg] 1.5 1Total Mass [kg] 18 15

4.6 Reaction Control System (Scott Weinberg)

The reaction control system is required to maintain 6 degree of freedom control. It is com-prised of 20 thrusters of two di�erent thrusts and 8 mounting locations. These locations werechosen to ensure proper control as well as to protect the astronauts and vehicle from plumeimpingement from the thrusters. A secondary concern was to prevent visual interference ofthe astronauts from the thruster nozzles. These thrusters maintain control over the worstcase scenario center of gravity shifts. Moreover, the RCS system can be used to safely landthe vehicle in the event of a main engine failure. Finally, all systems are designed to have twofault failure tolerance as per NASA requirements.

4.6.1 Control Requirements

The RCS system must maintain full, six degree of freedom control under the worst case centerof gravity shift. In the de�ned coordinate system, the vehicle will experience a maximum CGshift in the x direction of 0.16m, 0.00088m in the y direction, and 0.43m in the z direction.The downward (-z direction) thrust vector must remain within the CG of the vehicle. TheRCS must also be capable of safely landing the vehicle at maximum velocity and altitude. Thisrequires the RCS system to provide su�cient thrust to land in less than or equal to 2 earthg-forces and not use more fuel than the main propulsion system.

Center of Gravity Compensation

To maintain the control of the RCS system to all design parameters, the thrust vector of themain engine must always remain under the center of gravity. This means that the vehiclemust have su�cient control in pitch (moment about y axis) and roll (moment about x axis).Yaw (moment about the z axis) is not considered in this sense due to the fact that the mainthruster is in the x-y plane facing in the downward z direction. Thus, a CG shift in the zdirection does not a�ect the ability of the craft to yaw and would not set a constraint on thedesign.

To determine the necessary torques required to maintain control in the pitching and rollingdirections, the CG o�set was viewed as a moment arm. The force of the main thruster is40,000N. Since a moment is a force times a distance, a simple multiplication operation can�nd the required moments. Thus, the minimum required pitching and rolling moments werecalculated as follows:

NmmNDFM

NmmNDFM

xRoll

yPitch

35)00088.0()40000(

6400)16.0()40000(

=⋅=⋅=

=⋅=⋅=

52

Yawing was not constrained by center of gravity speci�c criteria. Thus it was assumed thatyawing torque should be su�cient to move the vehicle, but is not constrained by a minimumvalue as is the case with the pitching and rolling moments. The value of the yawing momentwill be determined after further analysis below.

Thrust requirements

There are 2 factors that determine the 3 axis thrusts that are required. The �rst factor thata�ects this design is the 2 earth g limit of acceleration in all directions. The second factor isthe ability of the vehicle to control itself and land in the event of a main engine failure. Thisis principally having su�cient downward (-z direction) thrust. There must also be adequatethrust in the x and y directions to traverse the vehicle. This will ensure the vehicle misseslarge obstacles when landing.

To stay under the 2.0 earth g limit, the burn time can be increased with a lower thrust. Theforce required of a thruster for two earth g's is calculated as follows:

ingB

MFn

Fsmgsmt

Vg ⋅=⇒⋅=⇒

⋅∆= 1

/81.9F/81.9

1

22

g2

(g=# of g's, ΔV=Velocity of vehicle, tB=Burn time, Fg=Force of deceleration, F=Thrustrequired for Fg deceleration, n= # of thrusters, Min=Inert mass of vehicle)

A ΔV=705m/s (half is used because this analysis just considers landing) yields a burn time of18.0 seconds for 2.0 earth g's. With an initial mass, Min=2380 kg, the force required is 46,600N. This is more than the thrust of the main engine. By dividing this thrust by the number ofthrusters, multiple engines could provide this force, greatly reducing the size and mass of thesystem. Moreover, if the g limit is set to 1.0 (36 sec burn time), then the force required is23300 N. The design can decrease the g's and divide the force between a number of thrustersto reduce mass and size of the system, while still maintaining the necessary control.

The x and y thrusts can be controlled with smaller thrust engines. The vehicle has fewconstraints in the x and y traversing directions. Only enough thrust is required to move thevehicle out of the way from obstacles before landing. This process can be started early, withthe start of the Lidar scan, further lowering the requirements of x and y direction thrust.

4.6.2 Thruster Selection

Burn Time Considerations

To determine the appropriate level of thrust for the RCS engines, burn time and number ofthrusters were considered. Below in table 13, burn time was listed in intervals of 15 seconds.The earth g force and thrust for a range of 1-8 thrusters was considered.

There were other considerations such as structural design, torque moments, mass of thethrusters, and mass �ow of the propellants that allowed for a more reasoned decision. More-over, as the structure of the vehicle developed there emerged 4 structural points on the truss'bottom that would be strong enough for the mounting of a thruster. In addition, throughcalculations, the necessity of booming out RCS thrusters to create su�cient moment armsfor torques emerged. For symmetry of control, this led to the conclusion of 8 downward (-z)pointing RCS thrusters that can be used for landing. However, all systems must be 2 faultfailure tolerant. As a result, if 2 thrusters fail, the remaining thrusters must still be capable

53

Table 13: Engine thrust and g-loads for varying burn times

Table #: Engine Thrust and g-Loads for Varying Burn Times Burn Time g's F 1 Eng F 2 Eng F 3 Eng F 4 Eng F 5 Eng F 6 Eng F 7 Eng F 8 Eng [sec] [-] [N] [N] [N] [N] [N] [N] [N] [N] 15 2.40 55944 27972 18648 13986 11189 9324 7992 6993 30 1.20 27972 13986 9324 6993 5594 4662 3996 3497 45 0.80 18648 9324 6216 4662 3730 3108 2664 2331 60 0.60 13986 6993 4662 3497 2797 2331 1998 1748 75 0.48 11189 5594 3730 2797 2238 1865 1598 1399 90 0.40 9324 4662 3108 2331 1865 1554 1332 1166 105 0.34 7992 3996 2664 1998 1598 1332 1142 999 120 0.30 6993 3497 2331 1748 1399 1166 999 874 135 0.27 6216 3108 2072 1554 1243 1036 888 777 150 0.24 5594 2797 1865 1399 1119 932 799 699

of landing the vehicle. Thus 8 downward thrusters will be installed, but only 6 will be neededto land the vehicle.

Mass constraints were based on the mass estimating relationship for the mass of a LOX/LH2small thruster: MLOX/LH2_Sm_Engr= 0.002395*T[N]

This value was used to compare di�erent burn times of a 6 thruster con�guration. 8 thrusterswere considered for mass, but only 6 thrusters were considered for thrust, burn time, and gload.

Table 14: Six engine thrust with equivalent 8 engine mass for varying burn times/ g-loadsTable #: 6 Engine Thrust with Equivalent 8 Engine Mass for Varying Burn Time/g-Load Burn Time g's F (6 Eng) M (8 Eng) [sec] [-] [N] [kg] 15 2.40 9324 179 30 1.20 4662 89 45 0.80 3108 60 60 0.60 2331 45 75 0.48 1865 36 90 0.40 1554 30 105 0.34 1332 26 120 0.30 1166 22 135 0.27 1036 20 150 0.24 932 18

54

0

20

40

60

80

100

120

140

160

180

200

0 20 40 60 80 100 120 140 160

Burn Time (sec)

Mas

s (k

g)

Figure 39: Burn Time vs. Mass of Thrusters

There is a point of diminishing returns at a 120 second burn time. Moving the burn time belowthis value will not demonstrate signi�cant decreases in mass for the thrust of 6 thrusters andthe mass of 8 thrusters. This result will be further constrained by the mass �ow of propellantsrequirement.

Mass Flow Considerations

It has been determined that an 8 thruster con�guration will be necessary. Two thrusters willnot be needed to accomplish the landing so the analysis will be conducted for 6 thrusters. Themass �ow of propellants is constrained to not exceed that of the main thruster.

In �gure 40, the possible thrust for a cluster of six and eight engines are compared on thebasis of mass �ow of propellant. The mass �ow of the main (40,000N) thruster is 10.2 kg/s.The nominal burn time of the main thruster is 20 seconds. This equates to a 1.7 kg/s mass�ow for the entire RCS system for a 120 second burn. This will ensure that the RCS systemdoes not utilize more propellant than the main thruster if it is needed to land the vehicle. Apropellant margin is added to the stores to accommodate attitude corrections and nominaloperation of the RCS. The mass �ows of the thrusters were calculated as follows:

gIV spe ⋅= eV

TM =

•

Here Ve=nozzle exit velocity, Isp=speci�c impulse (415sec), g=acceleration of gravity (9.81m/s2),=mass �ow of propellants, and T=thrust of engines. The thrust was varied for a six and eightengine con�guration to generate the �gure below.

55

1.7

1150853

0

0.5

1

1.5

2

2.5

3

3.5

0 500 1000 1500 2000

Thrust (N)

md

ot

(kg

/s)

6 Thusters

Main Engine

8 Thrusters

Figure 40: Thrust vs. Mass Flow for 6 and 8 thruster clusters

This plot shows that six thrusters with a 120 second burn should have 1150 N of thrust each.Eight thrusters with a 120 second burn should have 853 N of thrust each. However, since thissystem must be two fault failure tolerant, eight throttled thrusters will be used. They canthrottle down for eight thruster use, or in a contingency scenario where two thrusters fail, canbe operated at a maximum thrust of 1150 N to execute a landing. Thus the chosen design willinclude eight 1150 N thrusters placed to execute an emergency landing of the vehicle.

Attitude Control Thrusters

The vehicle has eight downward facing 1150N thrusters which can be used for attitude controland a contingency landing of the vehicle. These thrusters can provide z thrust, pitching, androlling moments. There must be additional thrusters to produce x and y lateral thrust as wellas a yawing moment.

There are no high level constraints on these degrees of freedom. A 450 N small size thrusterwas chosen to produce these forces and torque. They can be con�gured in such a way tocombine the laterally �ring thrusters to create the yawing moment. When designing thecon�gurations of the vehicle the amount and placement of these thrusters will be determinedto maximize control and minimize mass.

Nozzle Design

The nozzles of the two thrusters were designed to be a small as possible for two reasons.First is plume impingement. If the nozzles are short and squat (i.e. have a large expansionratio) the �ow will expand to safe pressures and temperatures very rapidly, making the chosenplacement locations safe for the astronauts. Second, and less important, is the aspect of visualimpairment of the nozzles. The lower in length the nozzles are, the less they will impede andastronauts' ability to see their surrounding area.

To accomplish both reduction in the diameter of the nozzle and length of the nozzle, chamberpressure was considered. As chamber pressure increased the nozzle length and diameterdecrease.

56

0.27

0.0670.00

0.20

0.40

0.60

0.80

1.00

1.20

0 500 1000 1500 2000

Chamber Pressure (KPa)

Nozz

le L

ength

(m

)

1150 N

450 N

Figure 41: Chamber Pressure vs. Nozzle length

0.21

0.0530.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

0 500 1000 1500 2000

Chamber Pressure (KPa)

No

zzle

Dia

met

er (

m)

1150 N

450 N

Figure 42: Chamber Pressure vs. Nozzle Diameter

From the above results a chamber pressure of 2 MPa was chosen. The limit that the systemcan provide is this value as well as the value of the chamber pressure for the main thruster.Thus, this pressure will reduce the size of the nozzles as well as simplify the piping and pressureregulation system as it can use the same pressurized lines as the main thruster.

57

Summary of Thrusters

Below is a chart of the two chosen thruster speci�cations:

Table 15: Thruster Speci�cationsTable #: Thruster Specifications Thrust Mass Isp Ve Mix Ratio mdot Po Ae/At Length Diameter [N] [kg] [sec] [m/s] [-] [kg/s] [Pa] [-] [m] [m] 450 1.08 415 4070 6 0.11 2.00E+06 115 0.067 0.053 1150 2.75 415 4070 6 0.28 2.00E+06 115 0.27 0.21

4.6.3 System Con�guration

To provide appropriate control the 1150 N and the 450 N thrusters must be placed in such away to minimize the number of thrusters (i.e. mass) while maintaining appropriate control.There are constraints such as the maximum CG shift that create a minimum for torques.These are a minimum of 6400 N-m for pitching moment (moment about y axis) and 60 N-mfor a rolling moment (moment about x axis). There is also a requirement, as discussed earlier,of eight 1150 N thrusters pointing in the downward z direction to ensure that the vehicle canland safely in the event of a main engine failure.

It was found that it was required to boom out four pods of thrusters to maintain necessarycontrol. This is particularly important to create the necessary pitching moment. A threepiece, space truss con�guration was utilized to boom the pods out to the appropriate distance.This will be discussed in greater detail later.

Below is a schematic of the chosen con�guration for the reaction control system:

Figure 43: RCS Con�guration

� All dimensions use shown origin.

58

� X is �length� Y is �width�

� Eight 450 N thrusters in x-y plane

� 2 each on pods 1,2,3,4

� Pods (±2.13, ±1.0) m

� Eight 1150 N thrusters in ±z direction

� 2 each on pods 1,2,3,4

� Pods (±2.13, ±1.0) m

� Four 1150 N thrusters in �z direction

� Mounted underneath vehicle

� Mounted (±1.7, ±1.45) m

This con�guration provides the necessary control for the vehicle. Below is the resultant controlforces and torques of the vehicle.

Table 16: Summary of RCS Thrust and Moment Control

Figure #: Summary of RCS Thrust and Moment Control

X Thrust Y Thrust Z Thrust Pitch

Moment Roll

Moment Yaw

Moment

[N] [N] [N] [N-m] [N-m] [N-m]

+ 900 900 4600 13710 5640 900

- 900 900 9200 13710 5640 900

Table 17: Roll Rates (Angular Acceleration)

Table #: Roll Rates (Angular Acceleration) Pitch Roll Yaw [deg/s2] [deg/s2] [deg/s2]

Full Vehicle 1279 183 29 Empty Vehicle 2663 527 83

Space Truss Boom Structure

It was necessary for the thruster pods to be boomed out a distance of 2.13 m in the x directionand over a distance of 1.0 m in the y direction. There are three mounting locations on thestructure to place the truss. Based on these two factors the lengths and angles of the spacetruss could be calculated.

59

The truss was constructed of Aluminum 6061-T6, which is the same material as the rest ofthe structural frame. The yield strength of this material is 310 MPa. The density of thisaluminum is 2700 kg/m3. A safety factor of 1.4 was considered. The elements of this trusswere designed to be resistant of axial (compressive and tensile), bending, and buckling failure.This resulted in the conclusion of rectangular beams of 0.0275m X 0.0275m. The results areas follows.

Figure 44: Boom Truss

Table 18: Boom Truss Structural Analysis

Table #: Boom Truss Structural Analysis Boom length b h Volume Stress σ Buckling Force Mass

[m] [m] [m] [m3] [N/m2] [N] [kg]

0.612 0.027

5 0.027

5 0.00046

3 273,320,00

0 15400 1.25

0

0.612 0.027

5 0.027

5 0.00046

3 273,320,00

0 15400 1.25

0

0.484 0.027

5 0.027

5 0.00036

6 216,084,00

0 24700 0.98

9

The maximum bending force was 1200 N which results in a stress of σv=71300 N/m2. Thisis well below the failure point of this material. Thus this design of the boom structure isresistant to the three types of failure.

4.6.4 Two-Fault Tolerance

The worst case, two failure situations for all six degrees of freedom was calculated. Below thereduced value of control is shown and it is proven that the vehicle still maintains necessarycontrol.

Moment worst case scenarios

� Two �Z direction thrusters on pods (1,2 or 3,4) failure

� Pitch moment drops from 13710 N-m to 8810 N-m

� Minimum required = 6400 N-m

� Two �Z direction thrusters mounted under vehicle failure

60

� Roll moment drops from 5640 N-m to 2300 N-m

� Minimum required = 60 N-m

� Two Y direction thrusters on pods (1,3 or 2,4) failure

� Yaw moment drops from 900 N-m to 450 N-m

� Minimum required = 0 N

Force worst case scenarios

� Two �Z direction thrusters (any of 8) failure

� -Z thrust drops from 9200 N to 6900 N


� Two X direction thrusters on pods (1,3 or 2,4)

� X thrust drops from 900 N to 450 N


� Two Y direction thrusters on pods (1,3 or 2,4)

� Y thrust drops from 900 N to 450 N


4.6.5 Piping and Valves

Piping and valves were necessary to provide proper pressure and propellant �ows to eachthruster. Each thruster required one two fault failure tolerant valve assembly. Each assemblyhas a mass of 0.5 kg and 20 are required. Piping has a mass of 0.5 kg/m length. To transportliquid oxygen and liquid hydrogen to each thruster or pod assembly requires 12.0 m of piping.

4.6.6 Summary of Systems

The reaction control system is made up of 20 thrusters. There are twelve, 1150 N thrusters andeight 450 N thrusters. The RCS maintains full 6 degree of freedom control through all �ightregimes. The system is capable of landing the vehicle in a contingency main engine failuresituation. Moreover, the system is designed to be two fault failure tolerant in all situations.Below is the mass breakdown for the entire reaction control system, which has a total mass of93.5 kg.

61

Table 19: Mass Breakdown for RCS

Table #: Mass Breakdown for RCS

Component Qty Mass/unit Total Mass

[-] [-] [kg/unit] [kg] 300 N Thruster 8 1.72 13.8 1150 N Thruster 12 2.75 33.0 Pod 1,2,3,4 Thrust Structure 4 4.18 16.7 Boom Structure 4 3.49 14.0 Piping 12.0 m 0.5 6.0 Valves 20 0.5 10.0 Total Mass = 93.5

5 Thermal (Amirhadi Ekrami)

5.1 Overview

The thermal control system of the Alshain lunar �ying vehicle will have the responsibilityof controlling and regulating the temperature of certain subsystems of the vehicle. Thesesubsystems include the seating structure of the astronauts, the avionics box holding the elec-tronics, the fuel cells, and the tanks storing the cryogenic fuels. Table 20 speci�es the safeoperating temperature ranges for these subsystems. In order to have a successful mission itis imperative for all systems to operate within their designated temperature ranges. Hence,the thermal control system will make sure these components operate safely during worst-casehot and worst-case cold scenarios. The worst-case hot scenario will be a daytime operationwith maximum internal power consumption along with full exposure to solar radiation. Theworst-case cold scenario will correspond to a mission under a complete shadow (possibly in adeep crater) during night operation with zero solar radiation exposure and minimum internalpower consumption.

Table 20: Operating Temperature Range

Subsystem Operating Temperature Range Seating Structure Under 320 K Electronics (Avionics Box) 250 K to 320 K Fuel Cells 330 K Cryogenic Tanks Liquid Hydrogen Liquid Oxygen

Under 20 K Under 90 K

62

5.2 Lunar Environmental Factors

The performance of the thermal control system will be heavily in�uenced by certain environ-mental factors. One of the important factors a�ecting the thermal control system will be theamount of heat �ux which is directed towards the system. These heat �uxes include internalpower consumed by the electronics and fuel cells, the heat emitted from the astronauts, directsolar radiation from the sun, solar albedo which is solar radiation re�ected o� of the lunarsurface, and also planetary radiation which is the radiation being emitted by the moon itself.Other important environmental factors a�ecting the system include lunar surface temperatureand ambient pressure. As will be explained later, the cryogenic tanks will be insulated usingmulti-layer insulation. An important factor for the operation of these layers is the ambientpressure, as they tend to drastically lose performance for pressures greater than10-3Torr dueto increased conductivity. So given the ambient pressure on the moon, the system will be ableto utilize multi-layer insulation. Table 21 lists a summary of environmental factors15 a�ectingthe thermal control system on the south pole of the moon.

Table 21: Lunar Environment Factors

Solar Irradiance 1367 (W/m2) Moon Albedo 0.11 Moon Emissivity 1 Minimum Temperature 50 K

Maximum Temperature 170 K Ambient Pressure 2 x 10-12 Torr

5.3 Seating Structure Heat Control

In order to keep the astronaut spacesuits safe from thermal damage, the seating structure ofthe vehicle must be maintained at a temperature below 320 K. For the analysis of the seatingstructure temperature, a worst-case hot scenario with a sink temperature of 170 K16 wasassumed, where the seating structure is 100% exposed to solar radiation. Additionally, thecontrol panels (50 W) and the astronauts (300 W) were assumed to be the only internal heatloads a�ecting this case due to their vicinity to the seating structure. The total surface areaof the seating structure, including steps and handles, was estimated to be 4.5m2. Given thatthe material of the structure is Aluminum 6061-T6 with an absorptivity of 0.2 and emissivityof 0.034, the equilibrium temperature reached on the seating structure amounts to 671 K.Therefore, in order to lower this temperature a vehicle coating had to be applied. For thismatter, the Aeroglaze A276 white paint was chosen. Aeroglaze A276 has the best performanceamong paints due to its high emissivity of 0.9 and low absorptivity of 0.23. Using this paint,in a worst-case hot scenario, the equilibrium temperature of the seating structure comes downto a safe temperature of 311 K. The radiation analysis for the seating structure (and allsubsequent sections involving heat radiation) is governed by:

15"Moon Fact Sheet." NASA. 1 Dec. 2008 <http://nssdc.gsfc.nasa.gov/planetary/factsheet/moonfact.html>.16"Thermal Control System for Exploration." NASA. 5 Mar. 2009 <http://www.nasa.gov/pdf/203096main_TEC%20Splinter-

Thermal%20control.pdf>.

63

5.4 Electronics and Fuel Cells Heat Control

5.4.1 Electronics Cooling

The electronics of the lunar �ying vehicle are an essential component of the vehicle and must bemaintained at safe temperatures during the entire duration of the mission. These electronicsfall into two primary categories: vehicle-mounted electronics (cameras, lights, star trackers,etc) and electronics placed inside the avionics box. Table 22 lists the electronics inside theavionics box along with their amount of power consumption. The power consumption of theseelectronics generates vast amounts of heat inside the avionics box. It is this heat that mustbe dissipated in order to maintain the electronics at a safe temperature.

Table 22: Avionics Box Electronics

Mass Data Storage 65 W

FPGA/DSP 75 W

IMU 45 W

WLAN 50 W

Flight Computers 100 W

S/Ka Band Transceivers 125 W

Interface Box 75 W

Total with 30% Margin 695 W

Initially, a passive radiator on top of the avionics box was considered for this purpose. However,the main problem with the passive radiator was the large amount of solar radiation coming intothe radiator; thus, requiring a greater radiation surface area to reach an acceptable equilibriumtemperature inside the avionics box. Due to the limited space available on the lunar �yingvehicle, a large radiator was not feasible to implement. To overcome this obstacle, opticalsolar re�ectors were chosen to be used as the radiation medium on top of the avionics box.By allowing the majority of solar radiation to be re�ected, while emitting heat from inside theavionics box, the optical solar re�ectors allow the radiation area to be minimized. Figure 23shows how a typical optical solar re�ector operates.

Figure 45: Optical Solar Re�ector

64

By using the optical solar re�ector, a minimum amount of solar radiation was absorbed, thuslowering the amount of overall heat needed to be dissipated, and thus decreasing the radiationsurface area to 2.3m2 and a safe equilibrium temperature of 306 K. Figure 46 shows how theradiation surface area di�ers between a passive radiator and the optical solar re�ector. Forthis analysis, the worst-case hot scenario was assumed with a sink temperature of 170 K anda maximum internal power consumption of 695 W.

Figure 46: Radiation Area Analysis

It is important to note that traditional optical solar re�ectors are composed of Quartz overSilver. However, due to the intense environmental conditions during landing and takeo� (dustand ejecta), these types of optical solar re�ectors are too fragile. Therefore, Silvered Te�on,know as �exible optical solar re�ectors17 were chosen to be implemented. Silvered Te�onoptical solar re�ectors give an emissivity of 0.87 and an absorptivity of 0.08518. In addition,due to their placement with respect to the vehicle, some percentage of solar radiation will beobstructed by the vehicle at all times, and will not reach them. Thus, a 70% view factor wasestimated for the radiation analysis.

5.4.2 Electronics Heating

As mentioned before the thermal control system must also accommodate for worst-case coldscenarios. As previously mentioned, during the worst-case cold scenario there will be no type ofsolar radiation. Parameters a�ecting this case will be the lunar sink temperature, which will beat 50 K, and also the amount of internal power consumption, which will be at a minimum. Theminimum internal power consumption for this vehicle will be during the 24 hour emergencysituation. During this time certain electronic components will not be operating. Table 23 liststhe electronics inside the avionics box which will remain running during a 24 hour emergencysituation.

17Gilmore, David G. Spacecraft Thermal Control Handbook. El Segundo, Calif: Aerospace P, 2002.18"Optical Solar Re�ectors." Qioptiq Space Technology. 21 Apr. 2009 <http://www.qioptiqspace.com/Data/Documents/Optical%20Solar%20Re�ectors.pdf>.

65

Table 23: Emergency Avionics Box Electronics

WLAN 50 W

Flight Computers 25 W

S/Ka Band Transceivers 50 W

Interface Box 25 W

Total with 30% Margin 195 W

Therefore, under these conditions, and given the emissivity, absorptivity, and 2.3m2 radiationsurface area of the optical solar re�ectors implemented in the previous section, the equilibriumtemperature inside the avionics box reaches 204 K. Certainly this temperature is below theminimum operating temperature allowed for the electronics. Therefore, additional adjustmentshad to be made to raise the temperature to a safe range. The �rst design considered was theuse of resistive heating inside the avionics box. However, this method would put extra load onthe vehicle's power source, and would not be feasible to implement for 24 hours. The seconddesign considered was the use of multi-layer insulation sheets to cover the radiation surfacearea during worst-case cold scenarios, in order to minimize heat emission. To accomplish this,however, the sheets had to either be manually deployed by the astronauts or be designed tobe deployed by a motorized system. In either case, the insulation sheets had to be storedsomewhere on the vehicle, when not in use. The chosen design, however, is similar to theinsulation covering concept, however instead of multi-layer insulation, a set of thermal louverswill be placed on top of the radiation surface area. The main advantage of thermal louversis that they will permanently be placed on top of the optical solar re�ectors and will operateautomatically when the need arises. Louvers are thermally activated shutters19 which requireno power to operate, and will open and close based on the temperature of the radiationsurface area. Therefore, when the temperature starts to drop in a worst-case cold scenario,the shutters will automatically close. The heat dissipation is then a function of the emissivityof the shutter surfaces in that closed position. Hence, given our surface area which is 2.3m2,to maintain a safe equilibrium temperature of 295 K, a surface with an emissivity of 0.2 isneeded. To accomplish this, aluminum paint is chosen as the surface �nish for the louvershutter surfaces. Figure 47 shows a typical thermal louver shutter system. It is importantto note that all other surface areas of the avionics box will be insulated using multi-layerinsulation.

Figure 47: Thermal Louvers

19"Thermal Control Louvers." Orbital Sciences Corporation. 5 Mar. 2009 <http://www.orbital.com/NewsInfo/Publications/Thermal_Louvers_Brochure.pdf>.

66

5.4.3 Fuel Cell Heat Control

Similar to the electronics inside the avionics box, the fuel cells will be placed inside an in-sulated case. Also similar to the avionics box, �exible optical solar re�ectors will be used tomaintain the fuel cells at their safe operating temperature during worst-case hot scenarios.Therefore, Silvered Te�on with an emissivity of 0.87 and an absorptivity of 0.085 will be uti-lized. During the worst-case hot scenario (Tsink = 170 K), two fuel cells will be operating andwill be generating a total amount of 635 W of heat. Therefore, to maintain the safe operatingtemperature of 330 K during this time, a radiation surface area of 1.4m2 will be needed.

During the worst-case cold scenario (Tsink = 50 K), one of the fuel cells will not be operating;thus, giving a total heat generation of 300 W. Given the emissivity, absorptivity, and the 1.4m2

radiation surface area, the equilibrium temperature reaches 260 K, which is obviously too low.Therefore, similar to the avionics box, thermal louvers will be used on top of the radiationsurface area to minimize heat radiation. Under these conditions, to maintain the temperatureat 330 K, the surface �nish of the louver shutter surfaces must have an emissivity of 0.3.

5.5 Cryogenic Tank Heat Control

The Alshain lunar �ying vehicle will be utilizing cryogenic propellants; namely, liquid hydrogenand liquid oxygen. Therefore, it is imperative to keep these propellants below their boilingtemperatures at all times. To accomplish this, multi-layer insulation will be used to insulatethe tanks and pipings and maintain the temperatures inside. As previously mentioned in theLunar Environmental Factors section, the moon's very low ambient pressure allows the use ofmulti-layer insulation. The multi-layer insulation will consist of Aluminized Kapton, giving ane�ective emissivity of 0.00220. According to the properties listed in Table 24 and the graphsshown in Figure 48, one layer of insulation will be enough for each tank. With one layer, thetotal boil-o� from each tank during a worst-case hot scenario duration of 32 hours (8 hourmission and 24 hour emergency) will be 0.04 kg for liquid hydrogen and 0.05 kg for liquidoxygen.

Table 24: Cryogenic Propellant Properties

5.6 Equator Design Considerations

The design requirements for the Alshain lunar �ying vehicle specify that the vehicle will solelybe used on the South Pole of the moon, and thus, must be designed around the environmentalfactors of the South Pole. However, it was of interest to the design team to see how thevehicle's systems would alter if the vehicle were to be used on the lunar equator. Namely, the

20Wertz, J. R., and Wiley J. Larson. Space Mission Analysis and Design. New York: Springer, 1999.

67

Figure 48: Multi Layer Insulation Analysis

thermal control system was of importance, because of the drastic increase in lunar temperatureat the equator. The main point of concern was the avionics box in a worst-case hot scenariowhere the lunar temperature reaches 400 K, along with 695 W of internal power consumption.However, according to Figure T.5, the optical solar re�ector's performance will greatly varyat the equator as compared to the South Pole. Using the same radiation surface area of 2.3m2,the equilibrium temperature inside the avionics box reaches 433 K, which is higher than themaximum allowable temperature. It is apparent in Figure 49 that even if the radiation surfacearea were to increase, the performance of the optical solar re�ector would not enhance, and asafe equilibrium temperature could not be reached by passive means of radiation.

Figure 49: Passive Radiation Comparison

An additional design considered for this scenario was the use of heat pipes. Heat pipes arevery e�cient heat rejection systems, and require no power to operate. They operate by thecapillary forces inside the pipe. However, heat pipes do not operate well under gravitationalforces because of the pressure imbalances that occur within the pipe. Given that the Alshainlunar �ying vehicle is designed to pitch and roll to certain angles, the heat rejection of the heatpipe will become a function of tilt angle; as the heat pipe tilts from a completely horizontalposition to an angled position, it will be exposed to the gravitational �eld of the moon andalso landing and takeo� forces. The heat load rejection by the heat pipe is governed by:

68

Analysis was done using several typical heat pipe designs21 with an e�ective length of 0.25 mand a working �uid of Ammonia. For an adiabatic temperature of 300 K, Figure 50 shows thatthe heat pipe will drastically lose its performance for even small angles of tilt. Given that theAlshain lunar �ying vehicle is expected to pitch and roll at angles of about 30 degrees, it wasconcluded that heat pipes could not be utilized on this vehicle. Therefore, to accommodatefor thermal control on the equator major changes need to be made to the thermal controlsystem. Namely, a pumping system must be used which requires more mass and power.

Figure 50: Tilt Angle E�ect

21Mills, Anthony F. Heat and Mass Transfer. Burr Ridge, Ill: Irwin, 1995.

69

6 Structures and Mechanics

6.1 Coordinate System (Edwin Fernandes)

The origin of the design for use in C.G. analysis was in the middle of the rectangular base (xand y coordinates). The origin of the z coordinate is at the bottom of the combustionchamber. The positive x-axis points in toward the front of the vehicle, the positive y-axispoints to the right of the vehicle (when facing forward), and the positive z-axis points up.The three axes are shown below.

Figure 51: Coordinate System

Figure 52: Coordinate System

6.2 Center of Gravity Analysis (Jarred Young)

In order to analyze the stability of the Alshain, the Center of Gravity would need to becalculated for the craft for both its pre-�ight and post-�ight con�gurations. However, as thisis a vehicle that is both in-�ight, as well as consistently expending fuel, the problem wouldhave to be at �rst bounded in order to understand the shifting of the CG of the Alshain LFV.

To analyze the CG, the following equation was used to calculate for pre/post-�ight con�gura-tions:

70

Table 25: Summary of CG Cases

For Pre-Flight con�guration, the standard loaded mass of each individual component on thecraft would be taken into consideration; the likewise can be said about the Post-Flight con-�guration, with the exception of the expendable items on the vehicle (such as propellant, forexample), as for a full mission, these items would either be depleted or o�oaded at the missionsite. Although time was not a main factor in the analysis, the two bounds of the pre/post-�ight considerations provide boundaries for CG-shift which would allow us to see the envelopewithin which the center of gravity acts.

Also, because our vehicle is a �ying vehicle, mass shifting is a huge concern. To this end, therewere 18 test cases in which to analyze our vehicle's shift. The cases are listed in table 25:

For all of the cases above, one of the most important limiting factors were the seating arrange-ments of the crew. In our structural con�guration, the crew is placed in a �stadium seating�situation where one crew member is sitting above and behind the crew member in the �rstseated position. The alternation of positions can a�ect the CG, especially when the two crewmembers are of di�ering masses (for example, cases 2 and 3).

The above table shows the co-ordinates and masses of all of the major components of theAlshain LFV. For such items as the crew, their seating positions are listed, but the massesare not listed so their respective masses can be taken into account for our alternating cases.For our propellant tanks, the mass of the tanks with no propellant are listed above. Theirrespective propellants and pressurants are listed by total mass on the table and are then splitequally into the number of tanks allotted for each of them above in the table.

71

Figure 53: Alshain Side Image

Table 26: Alshain Mass Balance

72

For each of the alternating cases in table 26, as well as the large amount of components toalternate in the preceding table, the CG Analysis was done in Microsoft Excel. For each case,the variables, such as crew mass and crew seating positions are accounted for, as well as theremaining factors of the CG Test cases. The following images are examples of one of the casesdone in Excel.

After the case is inserted into Excel, it is programmed to come up with the Pre-Flight CG,as well as Post-Flight CG. The transformation from one con�guration to the other involvesthe omission of some masses, such as propellant, payload (if o�oaded before or after �ight),etc. The graphs are also used to depict the shift of the CG from pre-�ight to post-�ightcon�guration for that case with respect to the Alshain origin, which is explain in our co-ordinate system section. Below is a summary of all of our CG locations, pre-�ight and post-�ight, for all of our case studies.

Comparing the numbers for each case, we can then evaluate the maximum CG shift in eachdimension of our vehicle in order to tell us the CG �envelope� of our vehicle, meaning thatwithin the boundaries set from the maximum shifts, our CG will never shift outside of thisenvelope (unless abnormally loading to due something outside of the Alshain mission pro�le).

6.3 Moment of Inertia Analysis (Jarred Young)

To further analyze the craft's stability during normal operations, the LFV's moments of inertianeed to be found. In order to do this, each component of the Mass Balance table needs to beaccounted for towards the Moment of Inertia calculations. From there, every component witha non-quadrilateral shape is approximated as either a sphere or cylinder and then the momentof inertia is calculated for each of these components. These moments of inertia are calculatedusing the following equations:

m=mass (kg)

r=radius (m)

73

Table 27: Sample Excel CG Case

74

Figure 54: Sample Excel CG Case

75

Table 28: CG Location Summary

Table 29: Max CG Shift Per Dimension

76

These moments of inertia are the centers of mass for each of these shapes. For all other shapes,the moments of inertia are calculated as follows:

x=X coordinate of position

y=Y coordinate of position

z=Z coordinate of position

These equations represent the moments of inertia of a point mass in three-dimensional space.For those components with a non-circular geometry, this is a fairly accurate approximation oftheir moments of inertia with respect to the axis of the vehicle. As this is the case, the co-ordinates on the Mass Balance table represent the respective center of mass of the componentsplaced on the Alshain.

To �nd the total moment of inertia, the Parallel Axis Theorem is applied to give the overallvalues of inertia for our craft. Here is the full equation used to analyze the Alshain Inertiamoments:

77

Table 30: Moments of Inertia Analysis

To compile all of these values easily, a MATLAB script was used to compile all of the co-ordinates of each component, their respective mass, as well as their respective geometries andradii (if applicable). For each of the geometries, a function �le was made with their respectiveinertia matrix including their component-wise moment of inertia formulae. Presented beloware the nominal moments of inertia.

6.4 Landing Gear (Andrew McLaren)

One of the primary critical structures for any �ight craft is the landing gear. Being able todissipate landing loads is critical to both the structural integrity of the rest of the craft aswell as the safety of the crew on board. While the Altair already has a presumably optimizedlanding gear assembly, the Alshain cannot borrow the design because the Altair's landing gearis for one use only; where the Alshain's landing gear must be reusable. For the mission criticalparameters, the Alshain's landing gear must be able to dissipate a 3 m/s vertical velocity anda 1 m/s horizontal velocity impact on landing. The Alshain must also be able to land on a 15

78

Figure 55: Number of Legs

degree slope and be able to negotiate a 30 cm tall obstacle anywhere within its landing zone.As with the rest of the LFV the landing gear will be designed to optimize mass.

6.4.1 Number of Legs

The �rst crossroad in determining the optimum leg design is to determine the number of legs.In order to constrain this decision we will assume that in the worst case loading scenario theAlshain will land on just one leg. This means that no matter the con�guration, each of thelegs must be able to take the entire landing load. In addition to this constraint the Alshainmust also not tip over upon landing. In the worst case tipping scenario the LFV will land ontwo adjacent legs. The amount of reaction moment generated by the center of gravity of thecraft to prevent tipping is a function of the leg length and the e�ective moment arm generatedas the projection of the length of the leg on to the plane of tipping. However since we areconstraining the size of the leg in order to take a given load, then the only variable is thenumber of legs. As is seen in Figure 55, for a leg of given length, r1, the e�ective momentarm, r2, is the perpendicular bisector of the triangle formed by the two legs that the LFV islanding on and the line between the two ends of the legs. Since we are optimizing for masswe want the longest possible ratio of r2 per weight of a system of legs of length r1. Since themass is given by a distinct leg size, the mass scales linearly with the number of legs. Thus forthree legs the ratio is cos(60)/3=.167, for four legs the ratio is cos(45)/4=.177, and for �velegs the ratio is cos(36)/5=.162. The maximum ratio is seen when the landing gear set up hasfour legs, thus our mass ideal design should consist of four legs.

6.4.2 Landing Load Acceleration

From a crew safety stand point our crew systems team recommended a landing load of no morethan 2 Earth g's. From the standpoint of trying to minimize the vehicle height this maximumsafe loading condition was chosen as the condition to design to. Since the landing gear on theAlshain must be reusable it cannot incorporate the crushable honeycomb interstitial from theAltair lander, thus linear springs were chosen as the energy absorption device. Solving fordynamic acceleration through the ordinary di�erential equation:

.5kx2 + .5mv2 = 0

In order to �nd the appropriate spring sti�ness we �rst solve this equation for the minimummass case, as that will be the case where acceleration is the greatest (13.2.4.1). Once thespring sti�ness is determined to be 53500 N/m it must be applied to the maximum mass casein order to determine stroke length (13.2.4.2). Now that the overall displacement, 0.77 m, is

79

Figure 56: Landing Tipping Acceleration

known it is necessary to split it into vertical (13.2.4.3) and horizontal (13.2.4.4) componentsin order to assess vertical height and tipping protection. The values for these components are0.73 m and 14 m/s^2 vertical stroke length and acceleration respectively, and 0.21 m and 6.4m/s^2 horizontal stroke length and acceleration.

6.4.3 Landing Tipping Acceleration

The next constraint of contention is to prevent tipping of the Alshain upon landing. Theworst case tipping scenario occurs when the LFV is landing on a 15 degree slope, with a 30cm obstacle under its upslope leg, moving 1 m/s horizontally in the down slope direction,and having no vertical velocity. Looking at the results of the horizontal acceleration analysisindicates that our horizontal acceleration worst case scenario is 6.4 m/s^2. Upon reviewingthe tipping analysis code, it becomes evident that continuing to handle the horizontal andvertical accelerations with just one spring would be infeasible. This can be easily seen byexamining the simple torque equation:

T = 6.4mLsin(15)− 1.6mLcos(15) = .11mL > 0

Where m is the mass of the vehicle, L is the distance from the down slope leg end to the centerof gravity, 6.4 is the acceleration, in m/s^2, induced by the spring, 1.6 is the gravity of themoon in m/s^2, and 15 indicates the angle, in degrees, the line L generates with the horizontalmoon surface when no obstacle is present (or the legs are signi�cantly longer than the heighteven with a 30 cm obstacle under the upslope legs). Figure 56 can assist in illustrating thissituation.

This equation shows that no matter how long the legs become, the net torque will always begreater than zero and thus the vehicle will always tip. Therefore separate springs are neededto handle horizontal and vertical acceleration. From this point we know that 6.4 m/s^2 is toomuch so we need to �nd an acceleration that will work. The code in 13.2.4.6 illustrates therelationship between a range of horizontal accelerations and the given moment arms that arerequired to prevent tipping at those accelerations. Note that these calculations are relying onutilizing the RCS to assist with counter-torque. Now that a range of accelerations is known,the code uses those accelerations to determine the corresponding spring sti�nesses, and othercode reads those sti�nesses to �nd the matching stroke lengths for the horizontal displacement.The optimum acceleration ultimately chosen, 0.95 m/s^2, was a function the chosen designconcept, which will be discussed later.

80

Since the horizontal acceleration decreased such a considerable amount, the net maximumacceleration seen was decreased, this allowed for a few iterations of increasing the verticalsti�ness and re-solving, arriving at the aforementioned �nal answer. As a result of this and afew other design changes along the way the �nal vertical stroke length became 0.82 m. Theremaining �nal sizing parameters include the minimum allowable height of 0.51 m (30 cmobstacle clearance +10 cm protrusion of rocket bell below the base of the structure +11 cmsafety margin so that the rocket bell doesn't seal with the ground and detonate), which stackson top of the vertical stroke to get a maximum uncompressed height of 1.33 m to the baseof the craft. The �nal length of the leg is 2.3 m total with a 1.64 m long footpad at the endof it that houses the horizontal deceleration spring. The footpad will be explained when thestructural design is discussed.

6.4.4 Ratcheting Spring Lock

Once the springs compress they have a large amount of stored potential energy that needs tobe dissipated. Classically dampers are used to gradually remove any excess energy howeverthe Alshain has forgone dampers in favor of ratcheting locks on the springs. There were twoprimary reasons for selecting ratchets over dampers. The �rst reason was stability. When theAlshain lands, if it rebounds too much, it may have a tendency to bounce in the lunar gravity,not something that is desired. The second reason was to have a limited auto leveling capabilitywhereby individual legs could be ratcheted open to compensate for uneven terrain. The onlypotential issue that could be thought of for ratcheting springs was how to release them when itwas time to take o� again. This is easily taken care of by having the capability to very slowlyratchet out the legs both before and during �ight back to the fully uncompressed position, soas not to generate any dynamic loads.

6.4.5 Cold Rated Spring (Bryan Han)

For the landing gear con�guration, spring will be used to put inside the legs to reduce the liftand landing forces. Also this spring will prevent from vehicle of tipping. In choosing the spring,there are some requirements. Since the spring will be working at the moon condition not anymaterial spring will work. One of the requirements is that it will function in the eternallydark areas of the moon, which the temperature is about 80 Kelvin. Therefore it should be amaterial that would not fracture in brittle manner at low temperature. So it needs to avoida ductile brittle transformation and so FCC material has to be used. Therefore the materialchosen for this requirement is copper-beryllium. Copper-beryllium does not undergo any lowtemperature allotropic phase transformations and work very well at low temperatures. Notonly that this spring has to work in the low temperature it will also has to work in fairly hightemperature of about 500 Kelvin, and this material can stand up to about 1300 Kevlin. Alsoit is found that this material is used widely in springs. Since this material spring will be takingforces during take-o� and landing, it needs to be sti� enough to not break. This material alsosolves that problem with having high strength and sti�ness. It has tensile strength of 240-280MPA, and modulus of elasticity of 125-130 GPA.

6.4.6 Leg Interface to Main Structure (Adam Halperin)

6.4.7 Leg Structural Design

In order to facilitate two springs and attempt to keep mass down by using as few members aspossible and by keeping to the philosophy that the shortest distance (and therefore lightestmass in this case) between two points is a straight line, the following design (Figure 57) waschosen. To handle the vertical acceleration torsion springs are placed at the attachment pointsof the legs to the base of the frame. Upon impact with the ground the legs splay along footpadrunners, both lowering the height of the vehicle and lengthening the width of the leg contactpoints out from the edge of the craft. The footpads also allow for the placement of a second set

81

Figure 57: Leg Structure Design

of linear springs that are relatively soft in order to provide the lower horizontal accelerationrequired to prevent tipping.

With tipping and the length parameter de�ned, loading can be analyzed. There are fourloading scenarios to consider, but of the four strictly compressive stress is far and away themost conservative case and does not contribute to the critical loading condition. The remainingthree loading conditions are Euler buckling, classical elastic buckling, and bending.

Euler: Fcrit = [(piL )2]EI

Classical: Scrit = Et

r√

3(1−nu2)

Bending: S = McI

In the Euler equation Fcrit is the critical failure load, L is the length of the column, E is thematerial's elastic modulus, and I is the moment of inertia. In the classical equation Scrit isthe failure stress, t is the wall thickness, r is the outer radius, and nu is Poisson's ratio. Inbending S is the stress experienced, M is the moment applied, and c is the distance from theneutral axis to the furthest surface of the cross section of the beam being bent.

Optimizing for all of these conditions also requires a manufacturability constraint that thethickness be at least 2 mm. After attempting to determine the optimized hollow cylinder massfor aluminum 6061 it was quickly discovered that in order to make the legs strong enough thelanding gear would be unacceptably heavy. Thus the switch to using the Titanium alloy wasnecessary (which Matweb has noted as being used in landing gear). Testing a hollow cylinderin the Titanium alloy proved a much more viable option. The result was certainly respectablehowever it was possible to take the optimization one step further. During the calculationsfor optimizing the mass for a hollow cylinder it was noted that the critical load scenario wasbending. While in bending a member must be its thickest at its root, but may be thinner at itstip and still not yield. So the calculations were repeated for a wide range of available radii andit was found that the limiting factor for the design radii was the machinability requirement for2mm thick walls. In order to be conservative the critical Euler buckling load was calculatedfor a cylinder of minimum radius and 2 mm thickness over the entire length of leg, explainingthe sudden change in radius pro�le trend. Having optimized using a tapered leg the actualleg assembly mass was brought to approximately 88 kg for all four legs combined. The footpad was modeled as a half cylinder that the end of the leg could rest in, resulting in each footpad having a mass of 5 kg. The support struts were 5 kg per leg, and the leg itself was justover 12 kg.

82

6.4.8 Leg Storage

The landing gear for the Alshain is too bulky to �t on the Altair lander in �ight con�guration.The gear will be unattached from the main body of the Alshain LFV and will be stored nextto it in order to facilitate space savings on the launch vehicle and lander payload bay.

6.4.9 Thermal Shielding (Bryan Han)

To shield the landing gear from the engine plume, landing gear will be covered with thermalshield. There are several kinds of thermal shield such as RCC, HRSI, FRCI, LRSI, and FIB.Thermal shield chosen for the design is FIB, which stands for Flexible Insulation Blanket. Thistype is chosen because of its low density, lightweight and ability to withstand high temperature.The temperature engine will produce is about 500 Kevin and FIB can hold up to 1700 Kevin.Comparing with all other types of thermal shield, FIB has the lowest density. Thermal shieldsuch as RCC, HRSI and FRSI have density in range of 192 to 1986 kg/m3, whereas FIB andLRSI has its density of 144 kg/m3. The reason FIB was chosen over LRSI is because FIBworks has higher temperature range and it requires less maintenance then LRSI. Also sinceFIB is made out of ceramic matrix composite, it is easy to be machined. This shield comes inthickness of 0.01 inches to 0.06 inches. When covering the landing gear legs with this thermalshield, the entire leg will not be covered, rather only half of it will be covered. It is becausesince the engine plume will only be contact with inner half radius of the leg, the outer halfof the leg will not experience the heat that requires thermal shield. The dimensions and themass of FIB are calculated in table 31:

Table 31: FIB Info�� Since the material has such a low density the total mass of covering four landing gear legsis only 0.087 kg. The chosen thickness is .01 inches (2.54E-04) because it gives the lightestmass. There is no problem with FIB getting ripped by engine plume, because this material isactually being used in spacecrafts that goes in and out of orbits.

6.5 Truss Structure

6.5.1 Design Rationale (Adam Halperin)

The primary loading cases, as described in the loads table, can be split into a stowed scenarioand a fueled scenario. In a stowed con�guration on the Ares-V launch vehicle we are subjectedto 6g axial (along z axis) loading, 2g lateral (x-y plane) loading, and random vibration loads.In a fueled scenario working on the moon we have 2g axial loading for landing and takeo�,along with a 2m/s^2 lateral load during landing. This causes the majority of the loading tobe axial, especially in the cases of functioning on the moon.

The con�guration of the components of the main vehicle is relatively short axially (1.46 m inthe z axis), while being relatively wide and deep (y and x directions respectively).

This causes large moments to be created by the primarily axial forces, while not allowing fora tall structure to distribute the loading.

83

Figure 58: Design Rationale Model

Table 32: Tube Inventory

Tube Inventory

Piece Designation

Number of Pieces

Applied Load (N)

Length (m)

Radius (m)

Thickness (m)

Mass (kg)

Engine Support

Engine Support Tube 4

17,000 (T) 1.51 0.016 0.0015 2.34

4 10,500 (C) 1.51 0.014 0.0015 2.04

Fuel Vertical LOX Support 16 2700 (C) 0.778 0.006 0.0015 1.66

Vertical LH2 Support 12 950 (C) 0.778 0.0045 0.0015 0.89

Horizontal Brace 4 1200 (C) 1.1 0.0055 0.002 0.67

Crew Vertical Support 4 1400 (C) 0.9 0.005 0.002 0.49

Main I-beam Crossbar 4 4950 (C) 1.56 0.01 0.002 1.91

Total Mass 10.00 �6.5.2 Structure Inventory (Bryan Han)

The inventory of structure beam supports is made. The calculations are all made withAluminum-6061 material; Elastic of Modulus of 68.9E9, passions ratio of .33, and yield stressof 276E6 is used in this calculation. For each structure supports length was determined.Therefore with its material properties and the length, radius was varied to obtain its beamdimension and mass, which can withstand designated load. For each structural member, stresslevel for tensile and compressive stress, Euler's buckling, and tubular buckling was calculated.Therefore it determined that at which case the tube would fail. When thickness and radiuswas found, thickness was rounded up to the nearest half millimeter for its machinability. Withthat machinability thickness, breaking load and its real mass was calculated, and below is thetube inventory:

This inventory shows with such applied load, length, radius and thickness, and all tubes totalmass. With such tube dimensions, breaking load was calculated and the margin of safety wasobtained in the below chart:

84

Table 33: Margin of Safety Table

Tube Inventory Piece Designation Applied Load (N) Breaking Load (N) MOS %

Engine Support Engine Support Tube 17,000 (T) 1.92E+12 1.10E+08

10500 (C) 13100 25

Fuel Vertical LOX Support 2700 (C) 3100 14

Vertical LH2 Support 950 (C) 1200 22

Horizontal Brace 1200 (C) 1350 12 Crew Vertical Support 1400 (C) 1450 2.5 Main I-beam Crossbar 4950 (C) 5200 5 �

Breaking load was calculated by using the following equations:

Euler Buckling:

Tubular Buckling:

Tensile/Compressive Force:

Also moment of inertia of tubes has the equation:

Then MOS: (Breaking load/Applied Load*Safety Factor) � 1

6.6 Crew Structure

6.6.1 Platform Design

For crew platform, honeycomb will be used to support the astronaut's seat. Each astronautweighs 170kgs, and at the impact of landing, astronaut will be experiencing force of two earthg's. To withhold that force which is 3400N, honeycomb need to be designed to withstandthat load. Since average person with spacesuit will take up about .75m for length and width.Stress is calculated using Force/Area, and came up with 1950 N/m2 . With safety factorof 1.4 stress is 2730 N/m^2. The honeycomb chosen was Aluminum Flex-Core honeycomb.It uses CR-PAA-5052/F40-.0013�-2.1� as a material, which it has speci�cation of phosphoricacid anodized coating, 5052 aluminum alloy, 40 nominal cell count in 12inch measured, .0013nominal foil gauge, and has density of 2.1lb/ft3. Therefore, for our design this honeycomb willhave dimensions of .75 width, .75 length, height of .25, and thickness of .0013 inches. Withthis dimension, this honeycomb can stand up to 1400000 N/m2. In calculating the mass thefollowing equation was used:

Volume of honeycomb:

85

a = apothem length

s = side length

n = number of sides

h = height

t = thickness

With this equation the volume came out to be 2.9E-4in3 and since in .75 by .75 areas there aretotal of 6 hexagons subtracting all the sides that overlap. Therefore multiplying the volumeby 6, the total volume of this honeycomb is .0017 4in3. So using given density property andcalculated volume, total mass of crew platform is 1.44 lb, which yields 0.66kg.

6.6.2 Roll Cage

6.7 Support Structure (Adam Halperin)

6.7.1 Support Structure Type

In order to counter these large moments, two design choices were considered. A single supportbase in the x-y plane and a truss structure were considered. However, the placement of thecrew and rocket nozzle created a restriction the placement of truss members. In order toadequately support the structure from bending under the large moments, truss memberswould have to be placed above, or in front of the crew members, which was unacceptable froma design standpoint. This led to the use of a single support structure, rather than a trussstructure.

6.7.2 Support Structure Placement

The placement of the support structure was ultimately chosen to minimize volume and mass,while not impinging on any component. The two options considered for placement of thesupport structure were around the middle of the fuel tanks, and �ush with the bottom of thefuel tanks.

Figure 59: Support Structure Placement

86

The placement at the middle would impinge on the crew access area, but would be placedcloser to the connection points for the fuel tanks, pressure tanks, and rockets. However, theplacement at the bottom would allow for a better attachment in a stowed con�guration withlanding gear removed, would allow for the components to be place closer together, and wouldhave a lower landing gear attachment point. The problem with the placement at the bottomwould be that all the components would require longer struts attaching them to the mainsupport structure, thus increasing the mass. The requirement of having an open crew accessarea, along with the various other bene�ts, led to the support structure placement �ush withthe bottom of the tanks.

6.7.3 Support Structure Shape

The shape of the support structure was determined by the x-y plane con�guration of thevehicle. The two sets of tanks and engine/crew area form three distinct sections, which wereoutlined by the support structure. This allowed for support struts to be placed around theoutside of each essential structure and have each main component supported in each direction.

Figure 60: Support Structure Shape

6.7.4 Beam Choice

The choices for beams to be used in the support structure were determined by an analysis ofinternal moments, shear forces, and torsion forces. Hollow tubular beams and I-beams wereselected for their strengths in taking these three types of forces. The I-beams are better thantubular beams in taking moment and shear forces, however, the tubular beams are signi�cantlybetter in withstanding torsion forces. This led to a trade study where the mass of tubularmembers was compared to I-beams and used to analyze the mass trade o� for torsion factors.The set of beams were analyzed for pure moment and shear forces, while ignoring the e�ects oftorsion. This analysis concluded that a support structure of torsion supported I-beams wouldbe 40 kg less massive than a set of tubular beams under the same moment and shear forceconditions.

This analysis concluded that the I-beams would be the more mass e�cient beam type if torsionprotection could be provided within a 40 kg envelope. However, if torsion e�ects became toolarge to be handled within the envelope, then tubular members would then become the moremass e�cient choice. The main source of torsion prevention in the structure was found to be

87

I-Beam vs. Tube

0

10

20

30

40

50

60

70

80

Lengthw ise Widthw ise Crew Support TotalM

ass

(kg)

I-Beam

Tube

Figure 61: I-Beam vs Tube Analysis

perpendicularly placed I-beams, which signi�cantly prevent torsional rotation. The e�ects oftorsion were large in the early iterations of development, but torsion conscious decisions onload distribution allowed for minimization of torsion e�ects in subsequent models.

6.7.5 I-beam Height Analysis

The height of the I-beam was a trade study done between mass, clearance for the rocketnozzle, height in stowed con�guration, and machinability. I-beams of 10cm height would be�ush with the, while anything larger would increase the stowed height of the vehicle linearly.I-beam masses varying between 7.5cm and 17.5cm in height were analyzed. The thicknesseswere calculated for each iteration individually in order to determine the optimal mass for eachheight.

I-Beam Design

0

10

20

30

40

50

60

70

0 2 4 6 8 10 12 14 16 18 20

Height (cm)

Mas

s (k

g)

Figure 62: I-Beam Height Analysis

The results of the trade study showed that mass savings signi�cantly drop after 12.5cm height,and after 15cm the mass begins to rise, due to the 2mm thickness requirement. A height of12.5cm was chosen for the I-beam support structure, as this allowed for 2.5cm clearance forthe rocket nozzle and is at the point of diminishing returns for mass savings.

88

Figure 63: IBeam Thickness Analysis

6.7.6 I-Beam Thickness Analysis

The analysis for I-beams was done by setting the total height and width of the I-beam cross-section and solving for the necessary �ange and web thicknesses.

The width of the I-beam cross section was chosen for each iteration of the design to keep the�ange thickness and web thickness relatively similar for stress concentration reasons, whilekeeping the size within reason for volume considerations. Throughout the analysis a minimumthickness for any given component was set to a minimum of 2mm for machinability purposes.For a given height the beams were analyzed for Euler buckling and Shear Force. The followingequations were solved for the thickness of the web.

Once the thickness of the web had been established, the thickness of the �ange was calculatedfor strength against the bending moment. This was done by using the bending momentequation and using algebra to solve for the thickness of the �ange.

6.7.7 Tubular Beam Thickness Analysis

In order to calculate the dimensions of a tubular beam for comparison to an I-beam, the outerradius of the tube was set half the height of the equivalent I-beam. This allowed for thecalculation of tubular thickness due to failure in both bending and shear cases. The formulafor bending failure was determined and algebraically solved for thickness.

89

Figure 64: Support Beam Loading Cases

The formula for shear force was similarly determined and solved for thickness:

Each tubular beam was analyzed for failure in both bending and shear force. The largerthickness of the two failure types was chosen as the minimum beam size requirement.

6.7.8 Support Beam Loading Cases

The support structure placement at the bottom of the vehicle allowed the assumption thatthe I-beam structure would be �rmly supported against the Altair cargo bay. This led to theassumption that launch loading would be transferred directly to the Altair cargo bay, makingthe 2g landing and take-o� scenarios on the Moon the worst case scenario for support beamloading. The worst case landing scenario considered was a two leg landing at 2g. A one leglanding is infeasible for static analysis as in any practical scenario the vehicle would eithertip, or fall into a multiple leg landing scenario. The 2g take-o� scenario has only one possiblyloading situation, with the exception of rocket or structure failure, which was assumed to bea critical failure anyways.

Load scenarios were developed for the various loading cases and used to calculate shear forcedistributions along the beams.

Landing on two legs:

� Fuel Tanks: 7.5kN per instance

� Crew Support: 5kN per instance

� Landing Gear: 20kN per instance

Take-o� and deceleration:

� Fuel Tanks: 3.75kN per instance

� Crew Support: 2.5kN per instance

� Rocket Force: 10kN per instance

Shear and moment diagrams were created for each considered loading scenario and used todetermine the maximum forces applied to each individual I-beam.

90

Figure 65: Beam Support Shear and Moment Diagrams

6.7.9 Support Structure Results

The three I-beam types were run through thickness calculations for web and �ange, using theirmaximum shear and moment forces. These calculations provided exact beam dimensions forthe optimal I-beam. These exact dimensions were used to determine the closest machinableI-beam.

Table 34: Support Structure Results

beam Web Thickness

(mm)

Flange

Thickness (mm)

Length (m) Mass (kg)

engthwise Exact

Actual

2

2

2.96

3

3.2 8.08

idthwise Exact

Actual

2

2

2.29

2.5

2.2 4.4

rew Support Exact

Actual

2

2

2.20

2.5

2.6 5.2

Once the actual I-beam dimensions had been set, the I-beams were tested for actual shear andmoment strength. The applied loads and breaking loads were then used in combination witha safety factor to determine the margin of safety.

MOS = (Breaking Load / Applied Load * Safety Factor) � 1

6.7.10 Truss System

Due to the choice of support structure placement, an array of small truss members werenecessary to join each speci�c component to the support structure. Each truss section wasarranged to avoid prevent bending moments through the load bearing members. To handlethe primarily axial, hollow tubes were chosen throughout the structure as the truss members.This member choice led to the analysis of each truss member for solely tensile and compressiveforces.

Tubular Member Under Axial Loading Analysis

Each tubular member was considered for four failure types: Tensile strength, compressivestrength, Euler buckling, and tubular �shell� buckling. The formula for tensile strength was

91

Table 35: Support Structure Margin of Safety

I-beam Shear Force (N) Moment Force

(N-m)

MOS (%)

Lengthwise Applied

Breaking

12500

50,000

4125

5795

0.35

Widthwise Applied

Breaking

3000

50,000

3300

4948

7.10

Crew Support Applied

Breaking

4670

50,000

3180

4948

11.14

the only formula for beams in tension. The formula was solved for thickness and used forbeam sizing with a given radius.

Tensile failure:

This formula, along with those for Euler buckling and tubular buckling, were used to determinethe thickness required for each failure case under compression. The largest thickness waschosen as the determining factor and the smaller thickness values were ignored.

Euler buckling:

Tubular Buckling:

In practice slight machining errors cause failures to occur at lower yield stresses. The as-sumption was made that machining errors would be no more than 15% of the wall thickness,corresponding to a failure stress at 50% of the ideal stress.

These formulas were applied to each beam design and the thickness for each beam was deter-mined by analyzing the thickness and mass characteristics of beams with varying radii. Theminimum allowable tubular thickness for tubular truss members was set to 2mm for machin-ability purposes. The results of the mass curve for varying radii was used to determine theideal dimension for any given truss member.

92

Figure 66: Tubular Member Mass vs Radius

Figure 67: Support Structure Crossbeams

Support Structure Crossbeams

Because of the perpendicular placement of the support structure beams, crossbeams werenecessary to maintain stability. The inner area has a natural protection against collapse fromthe rocket engine, but the outer sections require additional support. Two support beams wereplaced on either end of the support structure to prevent collapse of the support structure.

The worst case loading scenario for collapse of the support structure is 2m/s2 lateral forceduring a two leg landing, corresponding to a shear force of 4,000 N. This distributes evenly tothe symmetric crossbeams placed at 45 degree angles, leading to a compressive force of 3960N along the member. Running a tube analysis for compression over a range of radii leads toan ideal beam size of 1.6cm radius, 2mm thickness and 0.8kg mass.

Fuel Tank Support Structure

The fuel support structure is designed to connect the LOX and LH2 tanks to support structurebelow. The stowed (without fuel) load scenario and take-o� and landing scenarios (with fuel)

93

were considered. The empty LOX tank mass is 9.4 kg and the empty LH2 tank mass is 13.0kg, while the fueled LOX tank mass is 381.5 kg and the fueled LH2 tank mass is 63.5 kg. Therelatively large fuel mass causes the take-o� and landing scenarios to be the worst case forfailure. Under the worst case scenario the LOX tank creates an axial force of 7630 N and alateral force of 763 N. The LH2 tank, under the same conditions, creates a 1270 N axial forceand a 127 N lateral force.

94

Figure 68: Fuel Tank Support Side View

Figure 69: Fuel Tank Support Structure

This truss system involves a large portion of the vehicle mass and so it has a large e�ect onthe torsion of the I-beams. In order to minimize the torsion applied to the I-beams, the loadswere designed to be distributed to the corners of the support structure, and to the joint ofthe support crossbeams. This led to a system of perpendicularly placed sets of angled trussmembers.

A numbering convention was developed to designate the various sets of supports. Each set ofsupports is subjected to the same worst case loading scenario.

Each side refers to a pair of supports placed at a 45 degree angle as seen in �gure (refer tofront view of tank support). Each pair of supports takes one fourth of the axial and lateralforces from it's adjacent tank(s). This translates into compressive force within the beam byusing the equation.

95

Figure 70: Fuel Support Structure Diagram

These forces were used to calculate individual beam thicknesses for the fuel support system.

Rocket Support Truss

The rocket support truss is the only truss section that experiences signi�cant tension loading.The rocket support truss serves a dual purpose in that it supports the rocket, but also bracesthe central support structure against collapsing forces. Each support strut is set to reachthe nearest corners of the support structure in order to minimize the e�ects of torsion andalso to brace the structure against collapsing forces. The two considered loading scenarios forthe rocket support truss were the stowed con�guration with 6G axial and 2G lateral loads(compressive), as well as the take-o� case of 40kN (tensile). The tensile loading case duringtake-o� is calculated with a formula taking into account 4 beams oriented at an angle θ, whichis 53.3 degrees from the z axis.

This leads to an applied tensile force of 12.5 kN, but the compressive force must still beanalyzed. The compressive force in the stowed con�guration is calculated by a formula takinginto account a 6G axial force spread amongst all four support struts, with a 2G lateral forcetaken by two struts.

This leads to a compressive force of 3.757 kN, which is the most likely cause of failure, becauseit is a compressive force. The force of 3.757 kN and the length of 1.56m are used for the hollowtube analysis.

96

Figure 71: Full Skeleton Model

Results go here in the form of a brief MOS table once completed.

Fuel Tank Cradles

In order to secure the LOX and LH2 tanks within the structure, cradles were used. The bene�tof a cradle is that it does not create a hard point to the tank and act as a heat transducer.Semicircular cradles with rectangular cross sections were chosen for the large surface area forcontact against the tanks. Due to the di�culty of the analysis on semicircular cradles, thisanalysis was done in Pro-E FEA. A cradle cross section thickness (measured radially from tankcenter) of 1cm was chosen for machinability purposes. The cross section width (completingthe rectangular cross section) of the various cradles was varied under the di�erent loadingscenarios. Both LOX and LH2 tank cradles have 2G fueled axial loading as their worst casescenario, but the LOX cradle has 2m/s2 fueled lateral loading as its worst case scenario, whilethe LH2 cradle has 2G stowed lateral loading as its worst case.

Pic of FEA analysis done, along with results and MOS table

6.8 Vibration Analysis (Edwin Fernades, Joe Park)

Vibrational analysis of the Alshain was done using Pro/Engineer. The I-beam frame, tanksupport structures, RCS thruster supports, main thruster supports, and roll cage were allmodeled. The weight or the tanks, thrusters, and other components of the vehicles wereapplied to the model as point loads.

Shown above is a Skelton model of Alshain that was used for analysis in Pro/Engineer. Shownbelow is a table of the resonant frequencies of the �rst four modes of the Lunar Flying Vehicle.

97

Figure 72: First Mode Displacement

Figure 73: First Mode Displacement

Shown above is the displacement of the structure at the �rst mode. The �rst mode resonantfrequency of our vehicle came out to 6.784 Hz. The modal frequency requirement of a spaceshuttle payload is a minimum of 5.1 Hz. This means that using the Space Shuttle as anestimate, the stowed Alshain vehicle will not interfere with the stability of the launch vehicle.

6.8.1 Random Vibration Loading

Due to the similarity of having two solid rocket boosters, a conservative estimate of randomvibration loading was conducted using the Space Shuttle Orbital data22. Using the frequencyof 6.784 Hz found using Pro/Engineer and a damping ratio from the table below, a quasi-staticload was found using:

RLFn =√

πfnPSD4ξ

The resulting loads can be seen in the table below:

22Hap Ehlers. �Shuttle Orbiter / Cargo Standard Interfaces ICD 2-19001 Revision-L.�http://www.unitedspacealliance.com/icd/.

Table 36: Random Vibration Loading X-Y

Frequency (Hz) Power Spectral Density 20 - 90 .008 g2/hz 90 - 100 +9 dB/octave 100-300 .01 g2/hz

300 – 650 -9 dB/octave 650 - 2000 .001 g2/hz

98

Table 37: Random Vibration Loading Z

Frequency (Hz) Power Spectral Density 20 - 45 .0023 g2/hz

45 - 125 +9 dB/octave 125 - 300 .01 g2/hz 300 - 900 -9 dB/octave

900 - 2000 .001 g2/hz

Table 38: Damping Ratio vs Frequency

Table 39: Quasi-Static Vibration Loads

99

Figure 74: Alshain with Roll Cage

6.9 Roll Cage (Jarred Young)

One of the main tasks of the structure is to protect against all possibly predicted scenarios.This would include one possible scenario if the landing of the Alshain involved not clearing a30 cm obstacle (as per the Level 1 requirements). To this e�ect, a roll cage was required toprotect both the crew, as well as the vulnerable tank areas of the vehicle.

The tanks that contain the propellants and pressurant are most vulnerable when if the Alshainwere to roll over, as any obstacle that could puncture the tanks could cause a leak in a mission-critical system. In the propellant tank's case, we would lose vital fuel that we would need inorder complete the mission or return to base, as well as losing critical pressure within that tankto enable our pressure-fed system to operate. If the pressure tanks were to be punctured, therewould be a catastrophic failure for the vehicle. The pressure tanks are held under a pressureof 8.6 MPa; if the tanks were to be punctured, there would be an explosive force generatedwhich would severely damage the vehicle.

As a result, the roll cage was designed to protect against a rollover load (as dictated in Table40). The roll cage is a series of four arches mounted on the support structure of the vehicle,with two outer arches attached to the outer edges of the craft to protect the pressure tanksfrom rollover. These outer arches are connected by three crossbars 1 m in length to the innerarches to provide rigidity to the overall roll cage structure.

The inner roll cage is attached to the rear of the vehicle and extends to 30 centimeters aheadof the Alshain crew compartment and is attached to support spars. Through hardware testing,it was discovered that Alshain needed these extra support struts to provide clearance for theastronaut's feet and knees.

To design the roll cage, the pieces were approximated to be perfect circular arches made outof hollow tubes. The distance that the arch had to cover was set as the spread distance for thearch to model the forces. Here is a depiction of the equations used to analyze the roll cage,as well as a diagram.

100

Figure 75: Circular Arch Diagram

Table 40: Roll Cage Component Characteristics

These are the equations for Shear Force (V), Axial Force (F), and the Bending Moment of thearch. Using these with a circular cross-section for the beam, speci�cations could be set forthe beam that would optimize mass versus the stresses incurred on the beam due to Rolloverloading.

The roll cage numbers were established with a maximum force of 19 kN, which is the ap-proximate mass of the vehicle multiplied by the Rollover Loading (established from the LoadsTable) as well as factoring in a Safety Factor of 2. This allows for a conservative design toprotect against impact forces that may damage vital systems and crew, as well as still allowingfor a 30 cm clearance for obstacles as well as allowing the crew a 30 cm clearance to maneuverin the crew compartment.

101

The Crew Compartment member is a member that spans 3.48 m fore-to-aft of the Alshain andis 2.4 m high at its highest point; this design enables both crew members to have the properamount of clearance for their suits. It has a circular cross-section with an outer radius of 5cm and an inner radius of 4.75 cm.

The Tank Compartment member spans 2.14 m fore-to-aft of the Alshain and is 2.4 m high atits highest point. This allows the tanks to be clear from any obstacles that are de�ned by ourlevel 1 requirements. It has a circular cross-section with an outer radius of 3 cm and an innerradius of 2.85 cm.

These two member types are connected by their respective cross-bars to provide rigidity. Theyare designed as solid aluminum tubes with a radius of 1 cm. Although they are not tested forshear, their main purpose is to provide rigidity against forces in the y-direction.

The roll cage is primarily designed to take rolling in the forward x-direction, as CG Analysisindicates that the center of mass is more likely to shift forward during normal operation.However, if a roll were to occur in the y-direction of the Alshain, the cage could support thefull weight of the vehicle.

102

7 Crew Systems

7.1 Crew Placement Con�guration (Breanne McNerney)

Center of gravity (CG) analysis showed that arranging the crew stations one in front of theother created less CG shift than placing the crew members shoulder to shoulder at the front ofthe craft. The aft crew station was lofted to provide the second crew member with acceptablesightlines for contingency piloting.

7.1.1 Seating

The basic concept for the Alshain seating design was based on rover seating, because it isdesigned to accommodate EVA suited crew. Each seat includes two horizontal platforms thatvary in height by 0.343 meters and two vertical bars on either side of the top platform that areintended to support the astronaut's Personal Life Support System (PLSS). The entire seatingassembly consists of two seats aligned front to back with the back seat lofted 0.646 meters.

Speci�c dimensions for the seating design were based on an article testing rover seat ingressand egress in partial gravity environments. Table 41 shows the acceptable seating dimensionranges for the H-Suit (pictured in Figure X1) and I-Suit in 1G and 1/3 G.

Figure 76: Crew Seating

The 1 m wide platforms and 0.56 m spaced PLSS support bar dimensions were determined bythe geometry of the craft and PLSS respectively.

The angle of the PLSS support bars relative to the seating platform was chosen to be 90degrees, which is within the acceptable seat back angle ranges of both the H-Suit and I-Suit.The acceptable seat back angle ranges change little or not at all between 1G and 1/3 Genvironments, suggesting that these ranges are also applicable to the 1/6G environment.

103

Table 41: Rover Seat Dimensions

Figure 77: Seating Dimensions Front View

The seat height was chosen to be 0.343 m (13.5�). This dimension is the minimum acceptableseat height for both the H-Suit and I-Suit in a 1/3G environment. The minimum acceptableseat height value for the H-Suit increased from the 1G to the 1/3 G test; however, the minimumacceptable seat height value for the I-Suit decreased from the 1G to the 1/3 G test. Due tothese con�icting trends, it is di�cult to determine if 0.343 m will be an acceptable seat heightin a 1/6 G environment. However, this minimum value was selected in order to facilitateingress and egress because the seating platforms are also used as steps and therefore must beof an achievable step height for an EVA suited crew member.

The 0.41 m boot platform depth and 0.635 m seat platform depth were chosen to accommodatethe dimensions of a suited crew member as well as provide the necessary clearance for thecraft's main engine.

104

Figure 78: Seating Dimensions Side View

7.1.2 Sightlines

The forward crew member is the nominal pilot of the Alshain. The limitations on the forwardcrew member's �eld of vision is determined by the suit's helmet and visor assembly and lowerbody obstruction. The lower body obstruction only hinders sightlines in the bottom centerof the forward crew members �eld of view. The lower body reduces visibility in this area toapproximately 45 degrees inferior (greater than the 70 degrees helmet limitation).

The aft crew member is equipped with a set of control panels and has the ability to pilot thecraft in a contingency situation. However, the �eld of view of the aft crew member is morelimited than that of the forward astronaut. The view of the aft astronaut is further obstructedby the helmet and PLSS of the forward astronaut as well as the pressure tanks and roll cage.

The helmet and PLSS of the forward crew member obstruct the bottom center of the aftastronaut's �eld of view below approximately 30 degrees inferior. Tanks impinge slightly onthe inferior temporal region of the aft crew member's view �eld and due to the more recessedseating position, the roll cage structure creates an obstruction that is more central to the aftastronaut's �eld of vision than that of the forward astronaut.

7.1.3 Restraints

PLSS Restraint

The PLSS support bars incorporated in the seating design will also be used to restrain thePLSS. A repeating latch pattern will allow for custom positioning and a release bar will bepositioned within easy reach of the astronaut. A PLSS Engagement Con�rmation Light onthe control panel will signal to the crew member when the PLSS is properly engaged.

105

Figure 79: Helmet Sightlines

Figure 80: Lower Body Sightlines

Figure 81: Forward Crew Member Sightline Obstruction

106

Figure 82: PLSS Restraint Concept

PLSS latches are currently used to secure the PLSS during launch. These latches are designedto support the PLSS and suit in the event of a 9G crash loading. The combined mass of thePLSS and suit is approximately 70 kg. This results in a total load of 6180 N acting on thelatches supporting the PLSS. During a nominal mission, with a maximum acceleration of 2Gsand supporting a suited 95th percentile American male with a mass of 170 kg, the loadingon the PLSS restraints would be approximately 3300 N. Evaluated at the 9G shuttle crashloading scenario, supporting a 170 kg suited crew member would produce 15,000 N of loading.This additional loading could be accounted for by using multiple sets of latches to reduce theloading on each individual latch, or by a redesign involving stronger materials or larger loadingareas (cross-sections).

Boot Restraints

The 0.41 m boot platforms do not allow enough room for the aft crew member's boots tokick freely without contacting the PLSS of the forward crew member; therefore, they must berestrained. Boot restraints could not be placed directly on the boot platform without creatinga tripping hazard during ingress and egress. As a result, the boot restraints will be locatedon the back of the boot heel. This restraint con�guration will require a 90 degree knee bendangle which is less than the 120 degree maximum achievable knee bend angle (NASA STD3000).

Figure 83: Boot Restraint Concept

107

7.2 Crew Interface (Andrew Turner)

7.2.1 Overview

The Alshain LFV must be controlled and monitored at all times. In most scenarios, theseactions will be performed by an astronaut on board the vehicle. The astronaut will be ableto monitor all critical mission information as well as various non-critical tasks. The astronautwill be able to control these tasks through various input systems built into the craft and thespace suit.

7.2.2 Joysticks

The Alshain LFV will employ standard NASA joysticks on the craft. Each of the twoastronaut sitting positions on the craft will have an independent set of joysticks to controlthe craft, although only one set will be active at any given time. One joystick will be usedfor control of the translational motion of the Alshain while the other joystick will be used forcontrol of the rotational motion of the Alshain. Together, the joysticks allow the astronautto have six degrees-of-freedom control of the craft.

Figure 84: Left Joystick

Figure 85: Right Joystick

108

7.2.3 Controls Input

The astronaut will be able to interact with the Alshain vehicle by either using a keypad thatis attached to the right control stick or by utilizing a voice-command system in the helmet.

Keypad Inputs

The external keypad will be used for all mission-critical tasks on the Alshain. The keypadcontains 22 buttons including: 0-9 numerical keys, `+', `-`, A-F letter keys, clear, enter, reset,and proceed. The astronaut will be able to manually input any command using a combinationof the alphanumerical keys. The keypad will also have three covered switches for mission abort,engine cut-o�, and manual control. The mission abort switch will automatically send theAlshain back to its starting location upon activation. The engine cut-o� switch will manuallyshut o� the engine. It will only be used in an emergency when the craft is close to the groundso as not to cause a crash landing. The manual control switch will allow the astronaut tochange the level of user input during �ight.

Figure 86: Keypad

Voice Commands

The voice interface system will be able to give audio commands and feedback to the astronautthrough the headset as well as respond to vocal commands given by the astronaut. The voicesystem will only be used for non-critical tasks. The voice command system will be the primarycontrol interface for data displayed on the heads-up display (HUD). Similar systems have beentested on the ISS by NASA in cooperation with Xerox.

7.2.4 Displays

All necessary data and warnings will be displayed to the astronaut on-board the Alshainthrough a heads-up display and/or warning lights attached to the craft.

HUD

The astronaut will be able to monitor all necessary navigation and status data on the craftthrough a heads-up display (HUD). The HUD is not susceptible to glare issues and is alwayswithin the view of the astronaut so the astronaut will not have to turn away from a point ofinterest to view displayed information. Current work is being performed by NASA.

109

Figure 87: Voice Commands (Courtesy newscientist.com)

Figure 88: HUD (Courtesy Dr. William J. Clancy Desert-RATS Mission Simulation, 2006)

Warning Lights

Mission-critical warnings and status lights will be displayed on a light panel attached to theleft control stick. The panel will contain 8 lights and indicators including: master warning,radar fault, propellant level low, restart condition, navigation fault, program fault, PLSS latchengagement, and boot latch engagement. These warnings will also be shown on the HUD.

7.2.5 Control Panel Layout

Work Envelope

The control sticks must be placed within reach of both the 5th percentile American femaleand the 95th percentile American male. The NASA Crew and Thermal Systems Division atLyndon B. Johnson Space Center has performed tests with the advanced space suit designsthat will be incorporated into the Constellation architecture. These tests showed that theacceptable controller height was between 28� and 35�. The acceptable controller distance infront of the astronaut seat was determined to be from 7� to 14�. The Alshain was designed to�t in this range.

Sight Line Infringement

Tests performed with the advanced space suit designs concluded that the optimum positionsof the controllers would not infringe signi�cantly on the sight lines of the astronauts. Thewarning light panel and the keypad are placed within sight range of the astronaut on thecontrol sticks, but they are designed to interfere as little as possible with sight lines. Sightlines were the main driving force in determining which buttons needed to be included on thekeypad.

110

Figure 89: Warning Lights

Figure 90: NASA Astronaut Controls Test (Courtesy JSC2000E03273 Fig. 14)

Hardware

The control panels with a mock light panel and a mock keypad were used in the hardwarecomponent of the Alshain vehicle design. When the mock-up was tested by a suited person,the range and size of the buttons and lights was tested. Sight line infringement was alsotested in hardware. None of the controls interfered with ingress/egress of the àstronaut',and the àstronaut' was able to easily reach the control sticks. The hardware test resultedin an adjustment of the control panel however. One row of buttons was inaccessible by theàstronaut' and had to be moved further from the joystick.

7.2.6 Lighting (Pratik Davé)

The lighting con�guration on the lunar �ying vehicle must ensure su�cient illumination toallow safe �ight in daylight and night conditions. In all crew control areas, the lighting mustbe able to dim in order to accommodate di�erent contrast conditions. The lighting mustalso be able to properly function in an unpressurized, temperature-varying environment. For

111

Figure 91: Hardware Testing

this reason, halogen lamps were chosen over LED lamps, as LED lamps would require anextra pressurized housing in order to operate in lunar conditions. A summary of the lightingcon�guration can be seen later.

Vehicle Lighting

In accordance with NASA STD-3000, all lighting in crew control areas must be able to dimin order to accomodate di�erent contrast conditions. For the Alshain, this pertains to threeareas: the crew �ight control area, the crew ingress/egress area, and the cargo elevator area.Illumination of these areas will be provided by four 20 W white halogen lamps with a 40º�ood throw. Two of these lamps will be mounted above and behind the aft crewmember onthe two roll cage bars running on either side of the crew �ight control area, directed forwardin order to illuminate the control panels of both crewmembers. These two lamps will drawpower from the batteries in the event of a power contingency. A third lamp will be placedbelow the fore crewmember's foot platform and directed forward in order to illuminate theingress/egress ladder and surrounding area. The fourth lamp will be mounted in the centerof the roll cage crossbar located just above the cargo elevator. Each set of these lamps can bedimmed and controlled by knobs located in their respective areas. The lighting con�gurationsfor each of these areas are shown in Figure 92.

Surround Lighting

Four 50 W white halogen lamps with at least a 40º �ood throw will provide the illumination ofthe vehicle's surrounding vicinity. Each lamp will be placed on a corner of the vehicle's mainplatform. The lamps will be pointed towards the footpads of the landing gear legs in order toprovide su�cient illumination of each leg when landing. Once the Alshain has landed, theselamps can be used to light the surrounding vicinity of the vehicle. For this reason, these fourhalogen lamps must be dimmable in order to provide the correct amount of illumination forthe situation's contrast conditions.

The Alshain crew will be determining safe landing sites while the vehicle is in �ight; therefore,the potential landing sites will need to be properly illuminated from the vehicle. To assistwith this, two Polarion PF40 searchlights will be mounted on gimbals in the center of eitherside of the vehicle's main platform. Each of these lights operates at 40 W and has a range of1.5 km on earth. With the lack of atmosphere on the moon, it is possible that these lightswill operate at a higher range. The lighting con�gurations for each of these areas are shownin Figure 93.

112

Figure 92: Vehicle lighting con�gurations for (a) the crew �ight control area, (b) the crew ingress/egressarea, and (c) the cargo elevator area. One 20 W halogen lamp is located at each L.

113

Figure 93: Surround Lighting

Mission/ Contingency Lighting

When crewmembers are required to leave the vicinity of the Alshain in order to conductscienti�c missions or exploration during low-light conditions, portable illumination sourceswill be required. Each crewmember's suit will be standard equipped with either helmet-mounted or shoulder-mounted lights that will assist in these missions. If further illuminationis required, several �uorescent/incandescent combination lamps will be available on the vehicle.These lamps are portable and can be recharged using power from the vehicle if necessary. Eachlamp has a total life of 19 hours on a single charge (4 hours of �uorescent light and 15 hours ofincandescent light). Thus, these lamps can be used in contingency situations when necessary.

7.3 Debris Protection (Pratik Davé)

Comparing the Alshain to other lunar vehicles, the Alshain experiences some unique debrisprotection problems that have not been addressed in the past. This is because the Alshain isopen to the lunar environment and operating at much higher speeds than any lunar vehiclein the past. The Apollo Lunar Module and the upcoming Altair lunar module are entirelyenclosed and capable of being pressurized; shielding is integrated into the outer walls of thecapsule to protect from meteoroids and other debris. The same concept applies to designs fora pressurized rover. The Lunar Roving Vehicle (LRV), on the other hand, is not pressurized,and is open to the moon environment much like the Alshain will be. However, the di�erenceis that the LRV does not operate at as high a speed as the Alshain; the engine plume of theAlshain will eject lunar particles at much higher speeds than the wheels of the LRV. It ispossible that ejecta will be redirected back towards the Alshain, which poses a threat to thecrew. Proper debris protection is required for this reason.

7.3.1 Protection from lunar meteoroids

The �rst potential source for debris on the lunar surface is from space. The earth is surroundedby micrometeoroids and orbital debris and the moon has a similar surrounding environment.The di�erence is that micrometeoroids and orbital debris directed towards the earth burn up

114

Table 42: Lighting Con�guration Summary

upon entry due to the earth's atmosphere. The same does not apply to the moon. One canestimate the size of a meteoroid based on the probable frequency of a particle that size hittinga certain area using a lunar meteoroid �ux model. The �ux can be calculated using threepieces of information: the number of critical hits allowed during a mission, the area of thesurface that requires protection, and the duration of a mission.

durationmission

allowedhitscritical

tSA

nFlux

−

−−

×=

23

A critical hit is one that penetrates the shield; a critical hit in the case of the Alshain wouldbe dangerous to the crew, and so the propability of this occurance must be kept very low. Forthis analysis, one critical hit every 1000 missions has been chosen as allowable. The surfacethat requires protection in this case is the open surface area of the �ight control crew. The aftcrewmember is slightly shielded by the pressure tanks on either side and thus has a smallersurface area requiring protection (about 1.8m2), while the fore crewmember is entirely opento potential debris and thus has a larger surface area requiring protection (about 2.6m2).Finally, the nominal duration of a mission is 8 hours. The Alshain is designed to allow for 24hours of contingency on top of the nominal mission duration. For this analysis, the durationof a mission accounts for both nominal mission and contingency mission, 32 hours total (or32/365*24 years total).

yearm

hitsFlux fore ⋅

×== −

×2

1

24365

32101.1

))(6.2(

)001.0(

yearm

hitsFluxaft ⋅

×== −

×2

1

24365

32105.1

))(8.1(

)001.0(

23Akin, David. �The Space Environment.� ENAE483/788D, Lec. 7. Fall 2008.

115

Figure 94: Meteoroid Flux as a Function of Size

Thus, the �ux for the fore crewmember is nearly .11, and nearly .15 for the aft crewmember.Using the graph shown in Figure 94, these �ux values lead to particles from .02 cm to .03 cmin diameter. This value is the maximum particle size that must be protected against in orderto only experience one critical hit every 1000 missions. Taking the larger of the two particlesizes (0.3 mm in diameter), this is on the order of a rough grain of sand. Something this smallin size should be protected against by the shielding integrated into the astronauts' Thermaland Micrometeoroid Garment, the outer layers of the lunar EVA space suit that protectsagainst extreme temperatures and micrometeoroid penetration. Thus, shielding simply forthe purpose of protecting against normal lunar micrometeoroid activity is not necessary.

7.3.2 Protection during takeo� and landing

During the takeo� and landing periods of an Alshain �ight mission, there is potential thatrocks and dust from the lunar regolith could be ejected at very high speeds by the enginethrust plume (engine exhaust velocity = 3924 m/s). In the case that a piece of lunar ejectaencounters a nearby rock and is redirected back at the lunar �ying vehicle, the crew couldsustain serious damage; lunar ejecta could penetrate a space suit and breach the pressurizedbladder, or damage a vital life support system. Thus, the crew should be shielded from lunarejecta during takeo� and landing.

Because the Alshain is open to the lunar environment, adding structural shielding to the lunar�ying vehicle, such as aluminum or graphite epoxy plating attached to the roll cage or placedaround the crew �ight control area, would signi�cantly increase the vehicle's mass. Addingaluminum alloy plate shielding would increase the mass of the vehicle by at least 95 kg. Usinggraphite epoxy instead of aluminum alloy would still have a mass of at least 30 kg. Also,at least some parts of the structural shielding would need to be transparent in order to nothinder crew sightlines during �ight. Although transparent plastics are similar to graphiteepoxy in density, they are not similar in strength; thus, the portions of plastic shielding would

116

Figure 95: Conceptual image of TMG body debris shield

Table 43: Breakdown of TMG debris shield layers

need to be thicker than the graphite epoxy portions, adding more mass. To save mass onthe vehicle, a non-structural shielding con�guration was pursued. To further save mass, thisnon-structural shielding should be close to the astronauts' bodies, as this would lead to theshielding protecting the minimum surface area necessary. For these reasons, a debris shieldthat the crew could wear on top of a lunar EVA space suit was chosen.

Shielding against lunar ejecta is only required during the �ight portion of an Alshain mission.Wearing a debris shield when not in �ight (ie: during exploration or on experimentationmissions) would be additional, unnecessary mass for the crew to withstand. As a result, thisdebris shield should be designed so that it can be easily put on before takeo� and easily takeno� after landing. The debris shield has a mass of 4.6 kg and is designed as a sleeved blanketmade out of the same materials as the Thermal and Micrometeoroid Garment (TMG) layersof the lunar EVA space suit (see Figure 95 for depiction of debris shield). The shield will slideonto the arms of the crewmembers, cover all of their frontal and side areas, and clip to theoutside of the metal ring that connects the Hard Upper Torso to the space suit helmet. Thematerials, their areal densities, and the number of layers required can be seen in Table 43.

117

Table 44: Material Comparison - PMMA (Acrylic) vs. Polycarbonate

There are several advantages associated with using this style of debris protection. First,wearing it will double the amount of protective layers between the lunar environment and thespace suit's Liquid Cooling and Ventilation Garment layers. Wearing this TMG debris shieldwill also prevent any direct damage to the crew's space suits that could occur, as all directdamage will occur to the debris shield. The shield can be removed at any time, and can bereplaced or repaired at the lunar base if signi�cant damage has occurred. Another advantageof this TMG debris shield is that it can be a contingency radiation shield should there beradiation event while the Alshain is out on an EVA mission, as the materials of the TMG alsoprovide some protection from radiation.

The TMG debris shield cannot be used to protect the head of a crewmember, however, becauseit is not transparent and would thus interrupt the crew's sight during �ight. In order to protectthe head of a crewmember, a transparent plastic shield would be necessary. There are twopossible transparent plastics that could be used as a lunar ejecta debris shield: polymethylmethacrylate (PMMA, a.k.a. acrylic) and polycarbonate. A comparison of these materials'properties can be seen in Table 44.

Both of these materials are transparent, and depending on the grade and manufacturer, bothcan have the same tensile strength. The di�erence in the two materials lies in the impactstrength. Comparing these speci�c grades of plastics, the impact strength of polycarbonateis 13 times stronger than that of PMMA. Resistance to impact is a critical factor whenconsidering the protection of astronauts from high-speed lunar particles. For this reason,polycarbonate was the chosen material for the head shield.

The head shield has a mass of 7.8 kg and is designed to �t over and around the lunar EVAspace suit helmet and visor. As can be seen in Figure 96, the head shield will clip to twoitems: the PLSS and to the metal ring that connects the Hard Upper Torso to the helmet.Because the shield will result in additional mass that will be placed on the crewmember andthe suit, the head shield should only be put on when the crewmember's PLSS is locked intothe PLSS support structure on the vehicle. This will help transfer some of the mass to thestructure of the vehicle, and o� of the crewmember. For the same reason, the head shieldwould be removed prior to unlocking the PLSS support.

Wearing this head shield will increase the amount of protective layers between the lunarenvironment and the astronaut, and it will also prevent any direct damage to the crewmember'shelmet, as all direct damage will occur to the debris shields. The head shield can be removed atany time, and can be replaced or repaired at the lunar base if signi�cant damage has occurred.

118

Figure 96: Conceptual image of Polycarbonate head debris shield

Table 45: Summary of debris protection design

119

7.4 Contingency Procedures (Andrew Turner)

7.4.1 Extra Consumables

For the 24 hour contingency time period, extra consumables are needed by the astronauts.Each astronaut needs 0.85 kg of oxygen per day and 1.6 kg of drinking water per day. Eachastronaut also produces 1.0 kg of carbon dioxide that needs to be removed from the suit eachday. Each space suit requires 5.2 kg of water every 8 hours for cooling of the astronaut. Thiswater is evaporated into space during normal operation. The space suit also requires 120 wattsof power over a 24 hour period.

Figure 97: Human Needs (Courtesy Dr. David Akin's ENAE484 Life Support Lecture)

Drinking Water

Each astronaut requires 1.6 kg of drinking water each day. The space suit only has a smallpouch built in for the 8 hour nominal EVA. Each astronaut will be supplied with this extradrinking water through a feed tube directly into the helmet. There will be one tank of waterthat will supply both astronauts in an emergency. The tank will be a 0.2 meter diametersphere located underneath the front seat on the Alshain. The drinking water will be kept at20°C with insulation. The tank itself weighs approximately 0.37 kg.

Table 46: Drinking Water Tank Sizing

Total Drinking (kg) Drinking Density 3.7 998.2071

Drinking Volume Drinking Diameter

0.003706646 0.192044065

Cooling Water

Each astronaut needs a total of 15.7 kg of cooling water for a 24 contingency time period.With margin, a total of 33.3 kg of cooling water is needed for both astronauts. This wateris stored at 60°C so as to allow for possible re�lling by fuel cell output water. The tank is a0.15 meter long, 0.34 meter diameter cylinder with spherical end caps. This tank will also bestored underneath the front seat of the Alshain for ease of access and re�lling. The tank is a

120

cylinder to allow for stowage underneath the front seat. The tank itself weighs approximately3.3 kg.

Table 47: Coolant Water Tank Sizing

Water Tank Cylinders end caps (m^3) cylinder (m^3) total volume (m^3) diameter (meter)

0.0206345 0.013232337 0.033866836 0.34036 cylinder length (meters) 0.145509001

Oxygen

The oxygen used by the astronauts is supplied by the liquid oxygen oxidizer tanks of theAlshain. In a contingency mission there will be extra liquid oxygen within the tanks that willnot be used for power or propulsion. This extra oxygen will be drained o� the tank, heated upusing 2 watts per astronaut by the craft power system, and fed through a pressure regulatorto reduce the pressure from 348 psi to 4.3 psi. The �ow rate of the oxygen will be regulatedby the PLSS �ow regulators. This oxygen will be fed into the PLSS through an umbilical.

Carbon Dioxide Scrubbing

A total of 2 kg of carbon dioxide needs to be removed from the space suits to prevent poisoningof the astronauts. The air within the space suit will be cycled through the carbon dioxidescrubber using an umbilical. LiOH was chosen as the carbon dioxide scrubber for emergencysituations on the Alshain. LiOH is a common carbon dioxide scrubber. It has a density of1460 kg/m3. Approximately 5 kg of LiOH is needed to remove the 2.4 kg of carbon dioxideproduced by the astronauts with margin. The LiOH is stored in a 0.15 by 0.15 by 0.15 meterbox that will be stored underneath the front seat on the Alshain.

Table 48: Box Sizing

Mass Needed (kg) Cabon Dioxide Produced Margin Total Mass 5.04 2 0.4 5.04

LiOH Density (kg/m^3) Volume Needed (m^3) Box needed

1460 0.003452055 0.151132973

Power

Each space suit requires 120 watts of power for the period of 24 hours. This powers the fans,pumps, and communications within the suit. The PLSS batteries only last for 8 hours andare not rechargeable. This additional 240 total watts of power will be supplied by the Alshainpower system fuel cells or auxiliary batteries. The power will be fed into the PLSS throughan umbilical.

121

7.4.2 Radiation Contingency (Breanne McNerney)

In order to determine whether or not radiation shielding would be necessary during the 24 hourcontingency scenario, the probability of a harmful SPE and a mission failure both occuring ina 24 hour period was calculated. The probability of a large SPE is approximately 1-2 timesevery solar cycle (11 years) .

The probability of a large SPE occuring in a 24 hour period is approximately 0.05%. Thecalculated Alshain mission reliability is 99.2% (see section 12.2) therefore the probability offailure is 0.8%.

This gives a probability of both a large SPE and a mission failure occurring together a proba-bility of 0.0006%. The probability of this worst case scenario situation occurring at least onceduring a project lifetime of 250 �ights is equal to one minus the probability of no worst casescenario occurrences during the project lifetime.

The probability of one or more worst case scenario occurrences during the project life time isequal to 0.1% which satis�es the 1/1000 loss of crew requirement.

122

8 Hardware (Beal, Vasilak, McNerney, D'Amore, Sotak, Turner)

Hardware testing was an essential component to verifying the validity of certain systems withinthe design for the Alshain lunar �ying vehicle. Because of the complexity of crew interactionswith their surroundings and the limitation of CAD modeling software, certain parameterscould only be tested in a physical environment. These limitations led to the developmentof the hardware mock-up of the cockpit of the vehicle, including the roll cage, seats, controlpanels, and loading and unloading facilities.

8.1 Egress/Ingress (Sarah Beal)

The �rst piece of hardware was a method to load and unload payloads and incapacitatedastronauts to and from Alshain. Several di�erent methods were tested, including a winchsystems and an elevator. A ladder was also built to test the ability of astronauts to enter andexit the vehicle into the cockpit. Each of the scenarios required the ability to lift a downed170 kilogram 95th percentile male American Astronaut in 1/6g into the payload bay withminimum crew involvement to minimize the fatigue that would be associated with a high levelof activity. Because of the compressible landing gear, they must be able to lift payloads tothe maximum height of the landing gear. This occurs when the landing gear is uncompressed,with empty tanks and no payload, at 1.33 meters. The systems were also compared based onthe e�ect of lunar dust and corrosion.

8.1.1 Winch (Sarah Beal)

The �rst design that was considered was a winch system that would have the ability to liftthe payloads or incapacitated astronauts from the lunar surface and into the payload bay ofAlshain. The winch was designed to be operated either manually or automatically within thepayload bay itself and can be seen in �gure 98.

(a) Winch Sketch (b) Winch Model

Figure 98: Winch Concept

The winch is 1.1 meters high and 0.71 meters in reach, and has the ability to lift the payloadsin a locked position and then pivot 180 degrees around its base and drop the payload into thepayload bay. The system sweeps out an area of 0.26 square meters. It was designed at 1.1

123

meters high to sit at the waist of a 95th percentile American male Astronaut and the elbowof a 5th percentile American female Astronaut.

The hardware for the winch system consists of a winch that can lift 270 kilograms and has a0.15 meter handle which facilitates astronauts with a spacesuit glove. It also uses a pulley toguide the wire rope over the edge of the horizontal reach of the system. At the base of thewinch, a bearing is used to ensure the 180 degree mobility.

8.1.2 Elevator (Neal Vasilak)

Another feasible method for lifting payloads from the lunar surface to the main deck of thecraft is an elevator. Therefore, an elevator system was tested along with the winch.

Speci�cations

The hardware elevator consists of two aluminum guide tracks each 1.7 meters long spaced .9meters apart. The elevator platform is a 1 meter by .6 meter wooden board. Six guide wheelsare connected to the elevator platform, which allows it to ride up and down the guide tracks.Two wooden support arms are also connected to the platform for added stability.

Raising and lowering the elevator platform is a 50 in-lb torque, 3 RPM AC motor. The motoris attached to the underside of the main vehicle platform. This motor winds a nylon liftingrope that is attached to the elevator platform, which raises the platform up and down.

(a) Elevator Sketch (b) ElevatorModel

Figure 99: Elevator Concept

Advantages

An elevator has several advantages over a winch. The �rst is ease of use. An elevator can becontrolled using a simple push button control pad. Astronaut involvement is minimal.

Another advantage of an elevator over a winch is its smaller footprint. With an elevator, theelevator platform acts as the vehicles payload bay. Once a payload is loaded onto the elevator,the platform is raised to its �ight position and the payload remains there. With a winch, thereneeds to be room for the payload, room for the winch to be mounted and swing, and roomfor an astronaut to stand and operate the winch. This results in a 90% larger footprint for awinch system. This means a larger, more complicated, and heavier vehicle structure.

The elevator also has the advantage of being a symmetrical system. When the elevator ismounted on the rear of the vehicle it maintains its lateral symmetry, simplifying the center ofgravity calculations and minimizing use of the reaction control system.

124

Disadvantages

The mass of an elevator system is greater than that of a winch alone. However, the winchwould require greater vehicle structure mass to provide room for an astronaut to operate it,so this increase would be o�set.

Another disadvantage is that the elevator needs electrical power. This requires additionalfuel cells and batteries that a manually operated winch does not need. Requiring electricityand electric motors adds complexity to the system, adding more parts that could fail, anddecreasing its reliability.

Testing Results

The hardware elevator that was constructed worked successfully for lifting test payloads. How-ever, several �aws were noticed in the hardware design that needed to be corrected for theAlshain vehicle.

Figure 100: Elevator Testing

The �rst major �aw was the use of a rope to raise and lower the payload. A rope couldpotentially lose its tautness during accelerations and decelerations during �ght, causing thepayload to jerk around or the rope to unwind. The rope was also a single failure point, asit being cut or broken would cause the whole elevator to fail. Finally, the elevator platformcouldn't be raised completely to its stowed position, since the rope was mounted below themain platform of the vehicle.

The second major problem with the hardware elevator was that its tracks were mountedpermanently below the main deck level of the vehicle. This would place the tracks in theexhaust plume of the vehicle's main engine, causing them to melt during takeo�. The elevatorfor the Alshain vehicle had to be redesigned to be able to be stowed completely out of theexhaust plume and still be able to lower to the lunar surface level.

125

8.1.3 Chosen Design (Neal Vasilak)

After hardware testing was completed, the elevator was chosen as the payload loading system.The elevator was chosen due to its smaller footprint and symmetry. The smaller elevatorfootprint allows Alshain's stowed area to be nearly 10% smaller than it would be with awinch. This makes the vehicle easier to send to the moon as well as load and unload from theAltair.

However, the elevator needed to be redesigned after hardware testing was completed so thatthe elevator's tracks would no longer be impinged on by the vehicle's exhaust plume.

Instead of a platform riding up and down the tracks, the platform was redesigned to be rigidlyconnected to two 1.5 meter long aluminum rods. These two rods move up and down the insideof an aluminum tube, similar to the motion of a hydraulic cylinder. Each rod is geared, andis raised with an independent DC motor.

(a) Stowed Elevator (b) Deployed Elevator

Figure 101: Elevator Modes

8.2 Crew Stations (Breanne McNerney)

8.2.1 Ingress/Egress

In order to test the feasibility of the Alshain's seating con�guration, a mockup of the topplatform of the Alshain vehicle was constructed. Tests consisted of a suited test subjectascending a ladder, mounting the steps provided by the �rst seat and occupying the aft seat.The subject then descended the steps provided by the �rst seat, turned around, and occupiedthe forward seat. Finally, the subject turned around and climbed down the ladder.

Initial tests showed that additional hand holds would be necessary to facilitate ingress andegress as the subject needed to use control panels and PLSS restraint beams as support tocomfortably maneuver around the craft. The later addition of a rollcage provided structuralmembers on each side of the crew seating area that can be used for support during ingressand egress.

126

(a) Climbing Ladder (b) Ascending to Aft Seat

Figure 102: Hardware Ingress Testing

(a) Occupying Aft Seat (b) Dismounting AftSeat


8.2.2 Restraints

Gate latches were attached to both the PLSS restraint bars and the test subject's PLSS totest the feasibility of the PLSS restraint system. The subject required oral instructions inorder to successfully engage the latches, indicating that the crew member will require a visualof the latches in order to engage them properly (provided by mirrors or cameras) or that thelatching mechanism used must be designed to guide the PLSS into place.

8.2.3 Incapacitated Astronaut

Hardware testing was also run to determine the feasibility of intended rescue operations. Inorder to test the rescue of an incapacitated astronaut, a dummy was constructed to mimicthe weight distribution of a 95th percentile American male on the moon (1/6th the Earth'sgravity). The test subject was able to carry the dummy when handed it in a standing position,but was unable to lift the dummy from the ground. After handling the dummy, securingan incapacitated astronaut into the lowered payload lift seemed feasible to the test subject,however carrying the downed astronaut dummy up the ladder and securing it into one of thecrew seats did not. Future testing will include securing the dummy into the payload lift. Theresults of the incapacitated astronaut testing showed that a one person rescue mission maynot be feasible or may require special tools to assist with the rescue.

127

(a) Occupying FrontSeat

(b) Descending Ladder


Figure 105: PLSS Restraints Hardware Testing

9 Avionics

9.1 General Overview (Kush Patel)

Alshain's avionics system controls most of the vehicle's onboard systems. Its functions includeautomatic determination of the vehicle's status; operational readiness; performance monitor-ing; digital data processing; communications; guidance, navigation, and control.

9.1.1 Requirements (Kush Patel and Nick D'Amore)

The design of the avionics systems was guided by the following level one requirements forAlshain. These requirements must be met for a successful design and implementation of thevehicle:

� LFV shall be capable of autonomous �ight and landing at a planned base landing site

� LFV shall be designed for crew-guided stable �ight in all mission phases

� LFV shall be capable of being controlled directly, in teleoperation, and autonomously

� LFV shall be capable of communicating at HDTV rates direct to Earth

128

Figure 106: Incapacitated Astronaut Testing

� All critical systems shall be two-fault tolerant, with instrumentation for status monitoring

The �rst of these requirements dictates that, during nominal operations, the vehicle must havesu�cient sensing capabilities to identify any characteristics of the intended landing site thatcould pose a hazard. This includes rocks and other surface features 30 cm or larger as wellas slopes 15 degrees or greater because these are the maximum hazards that the landing gearhas been designed to encounter.

The second and third of these requirements dictate that allowances be made for several dif-ferent levels of human involvement in vehicle guidance. This has implications for the designof the guidance, navigation, and control system, as well as for the crew interfaces discussed inanother section.

The fourth of these requirements does not de�ne an explicit data rate. The rate necessary totransmit HDTV video depends upon the resolution, frame rate, and compression algorithmemployed. For the purpose of designing Alshain, the required data rate for nominal operationswas set at 100 megabits per second (mbps) to allow for transmission of high quality HDTVvideo or the equivalent amount of data. In the event of a fault, under which circumstancesdiminished mission parameters are acceptable, 30 mbps was de�ned as the minimum data ratethat would provide acceptable HDTV performance.

The �fth requirement dictates that Alshain's avionic systems must be designed to withstandmultiple failures through redundant hardware and software. These are managed by a complexof four computers to ensure robustness to two computer failures. After one failure in a system,redundancy management allows the vehicle to continue on its mission. After a second failure,the vehicle still is capable of safely returning to a pre-prepared landing site at the base.

9.2 External dependencies (Kush Patel)

9.2.1 NASA Lunar Surface Architecture

As part of the Constellation program, Lunar Relay Satellites (LRS) will be put in lunar orbitfor communications and navigation. As communications relays, these satellites will serve thesame purpose that TDRS serves for near earth network communications.

129

Overview

Alshain will plan its missions such that a LRS will be overhead for the duration of the mission.These satellites will primarily be used for command and telemetry exchange between thevehicle and the lunar base. However if Alshain is at a location where it can not make direct toearth communications link to the Deep Space Network (DSN) antennas on Earth, LRS' willbe used as a medium to creak the link.

Speci�c dependencies

The orbits of the LRS' will be very important for Alshain's missions. The amount of time thatone LRS is in line of sight to the vehicle and the lunar base will help determine the time frameof the mission. Alshain will also need to know ahead of time the scheduling of LRS operationsas other Constellation missions might also need to make use of these LRS'.

Lunar Relay Satellites

One of the reasons for the LRS' as mentions above is to support communications for theConstellation program. LRS' will be in 12 hour orbits around the moon with eight hoursexposure over the south polar region. These satellites are capable of one-and-two way rangingand are fully capable of relaying both S and Ka band communications. They also have theability to store and forward data with 300 gigabyte capacity.

Lunar Communication Terminals

Alshain is required to communicate directly to the lunar base. Lunar Communications Termi-nals will be mainly used to server this purpose. These terminals will be erected very close tothe lunar base and are capable of transmitting both S and Ka band along with both 802.16eand 802.11. They are also capable of one way ranging for emergency return back to the base.

9.3 Communications (Jolyon Zook)

9.3.1 Frequencies Speci�cations

NASA has speci�ed Ka and S-band channels for the constellation architecture, including bothbetween a potential lunar rover and a orbiting lunar relay satellite (LRS). A direct-to-Earthlink has not been speci�ed in Constellation literature and it is assumed that this would be ableto use the same frequency as the LRS back link with an Earth-bound station (not a TDRSSsatellite).

Ka band channels

A Ka-band link will be established to transmit and receive high-bandwidth data such as video,navigation or scienti�c data. It will be gimbaled such that it is capable of attaining pointingaccuracy necessary for communication between both a Deep Space Network (DSN) 34 mground station and an LRS orbiting at a maximum altitude.

S band channels

An S-band link will carry time-sensitive and safety-critical data such as voice communica-tions, tracking telemetry and control (TT&C), and vehicle status monitoring. Total S-bandbandwidth will be about 250 kbps.

130

WLAN

The Alshain vehicle will be an integral part of the planned wireless local area network (WLAN)on the lunar surface and as such will have the capability to use commercial protocols such as802.11 and 802.16e. The speci�c protocol to be used has yet to be chosen by NASA.

9.3.2 Architectural Model

Physical Layer

As already mentioned, Ka- and S- bands are planned for use between Alshain and Earthand between Alshain and other surface nodes, including the lunar base. These links will useright-hand circular polarization in order to retain compatibility with the lunar reconnaissanceorbiter (LRO) and also in case the relay satellites are spin-stabilized. While the LRO usesquadrature phase-shift keying (QPSK) modulation, we anticipate that by 2012 space-ratedsoftware-de�ned radio (SDR) systems will be available such that a lunar assets will not bebound to one particular modulation technique.

Infrastructure

The Constellation architecture calls for 8-hour coverage for every 12 hours in the polar regionsbeing considered for exploration. This means that at least one LRS will be available fornavigation and data relay during these times. As alluded to above, the LRO satellite maybe capable of providing surface mapping from its lunar orbiter laser altimeter (LOLA) orother instruments, provided it has radio coverage or the region in which Alshain will beoperating. In addition, a lunar communications terminal will serve as a WLAN hub and as adata relay between other surface nodes and either an LRS or Earth. It will also have routing,MUX/DEMUX, and ranging capabilities. Several of these terminals may be deployed.

9.3.3 Antennas

HGA

Parabolic dish and phased array designs were considered for transmitting high-bandwidth mis-sion data. Phased arrays were discounted for several reasons. First, the primary advantage ofphased-array systems is their potential to be electrically steered, enabling in-�ight transmis-sion without the need for gimballing. However the power and computational load, in additionto the extra mass of such a system with current technology, make it impractical for our pur-poses. Moreover, there is currently no requirement for high-bandwidth transmission in-�ight.However if such a need were to arise, and the technology had advanced, a phased-array systemshould be reconsidered.

The parabolic dish selected will be gimballed such that it has su�cient pointing accuracy toestablish a reliable link with the DSN 34 m dishes.

Required pointing accuracy can be determined by the characteristic half-power beam width,which can be calculated as: HPBW = k*l/D where l is wavelength, D is antenna diameter,and k is a constant accounting for the antenna shape, for parabolic dishes around 70 degrees.Therefore the half-power beam width for our design is about 1.2 deg. To maintain su�cientlink margin, the pointing accuracy should fall within this range. Current two-axis space-ratedgimballing systems have achieved accuracies on the order of 50 arcseconds, well within therequired accuracy.

131

LGA

We required a low-gain, omnidirectional antenna to transmit and receive critical data whilein �ight since reliable pointing will not be available during this time. Such an antenna systemmust have su�cient gain in both the azimuth and elevation directions to establish reliablelinks other ground nodes, the lunar relay satellites, and the DSN dishes on Earth.

The vehicle will have two LGA's: one mounted on the landing gear facing down to prvidecoverage of the ground while in �ight and one on top (above the level of astronauts' heads soas not to expose them or their radios to signi�cant amounts of radiation) providing coverageof satellites and Earth.

9.4 Command and Data Handling

The Command and Data Handling (C&DH) system behaves as LFV's brain. It distributescommands, records telemetry, and keeps various components' status updated in real-time.

The C&DH is comprised of an enclosure, a backplane, four single board computers, S-bandand Ka-band communications interface boards, data storage boards, a housekeeping and digitalinput/output board, and analog data acquisition boards. All electrical connections betweenthese components are made via the backplane for internal power distribution and PCI bus datatransfers. The interfaces between the C&DH and other avionics components are connectedthrough a SpaceWire network. Network provides for a standard that enables high and lowdata rate communication between avionics components over the vehicle. This network highlyincreases speed and reliability in space-�ight systems.

9.4.1 Equipment (Kush Patel)

Mass budget

The following presented mass estimates for the avionics components come from publishedcredible research papers and speci�c company technology technical documents.

The Lunar Reconnaissance Orbiter (LRO) has very similar command and data handling re-quirements to Alshain, though without the strict redundancy requirements of Alshain. TwoC&DH boxes of this size, including all subassemblies, total approximately 41 kg. However,Alshain will require some additional components not present on LRO. Unfortunately speci�cmass �gures at this level were not available, so the additional mass had to be estimated fromother sources. Taking a look at the general weights of circuit cards, it was estimated thatthe vehicle would need an additional 17 kg of components, bringing the total avionics massestimate to 119 kg.

Volume

The volume estimates of the �rst four components in the table have come from publishedresearch and/or speci�c company technical documents that are referenced at the end of thisreport.

The volume estimates of one C&DH box including all of its subassemblies comes mainly fromNASA's LRO mission. The speci�cs of the LRO mission is very similar to what Alshain'srequirements are. So for this level in design, it is best to research a bit more in to theiravionics box and go with this estimate.

132

Table 49: Avionics Mass Budget

Components: # of units Total Mass (kg) LIDAR 1 11 Radar 2 19 Star Trackers 2 2 Video Cameras 2 3 Inertial Measurement Unit 3 6 Parabolic Antennas/Amplifiers 2 12 Status Monitoring Instruments 5 Box/Casing/Wiring 58 Computers / Circuit Cards 4 Housekeeping Input/Output Card S& Ka-Band Communications Card Multifunction Analog Card / Backplane Table 50: Avionics Volume Budget

Components Dimensions Star trackers .178 m (L) .203 m (W) .102 m (H) Video cameras .203 m (L) .127 m (W) .076 m (H) Inertial Measurement Unit .203 m (L) .152 m (W) .152 m (H) Parabolic Antennas/Amplifiers .66 m diameter Box/Casing

Single Board Computer / Data Storage Card Housekeeping Input/Output Card S & Ka-Band Communications Card Multifunction Analog Card

Backplane

.5 m (L) .3 m (W) .3 m (H)

9.5 Guidance (Nick D'Amore)

9.5.1 Overview of Flight Plan

The nominal �ight plan for the Alshain is a modi�ed ballistic trajectory. The vehicle will beginwith a vertical climb of approximately 10 meters to clear any nearby terrain features beforebeginning to pitch toward the target (approximately 45 degrees depending upon altitudechange) and increasing engine thrust to begin the acceleration phase of �ight. Once theacceleration burn is complete, the vehicle will pitch back to a level orientation to enableproper pointing of communications and navigational equipment. The vehicle will then remainin a coast until it approaches the target. It will then pitch back to begin the deceleration burnand transition into a propulsive glide to enable the crew to visually inspect the landing siteduring the �nal approach.

9.5.2 Landing Site Selection and Hazard Avoidance

Prior to launch, the vehicle shall have a nominal target set in the computer system. Once thevehicle comes into range of the intended landing area, a lidar scan will be conducted for �nallanding site selection. In autonomous mode, the computer will automatically identify a sitesu�ciently free of hazards for a safe landing to be accomplished. The crew, if present, willhave the option of overriding this by designating an alternate landing site or switching into a

133

manual control mode.

9.5.3 Control Modes

The vehicle will have control modes for three di�erent levels of human involvement during�ight.

Autonomous

The default mode of operation will be autonomous control, in which the onboard computersystem manages all aspects of �ight without the need for any external intervention. Thedestination will be speci�ed prior to launch either by the crew or from a remote location via thecommunications system. Onboard software will then determine a suitable trajectory for �ightand display this trajectory information to the crew (if present) in addition to transmittingit to mission control. The vehicle will then require a proceed command, delivered eithervia the crew interface or remotely, before initiating launch. Automatic feedback loops willmaintain the desired trajectory. At an altitude of 100 to 200 meters, the vehicle will completethe deceleration burn, transitioning to a propulsive glide. A lidar scan will be conducted toassess the suitability of the targeted landing site, redesignating the target if necessary to avoidobstacles. If desired, the crew may override this target selection and specify an alternatetouchdown site.

Direct

The crew shall have the option of switching the vehicle into a direct control mode in whichautomatic control loops will maintain pilot-speci�ed rates, both in translation and rotation.These desired rates will be speci�ed via the crew interface described in the Crew Systemsportion of this report. This mode is intended for short range hover and glide operationsduring which the pilot may prefer to control the vehicle directly rather than simply specifyingthe �nal target as is done in the autonomous mode.

Teleoperation

During teleoperation, �ight will proceed in the same manner as in the autonomous case withthe exception that the lidar scan will be transmitted to a remote pilot and the vehicle willwait (transitioning into a hover if necessary) until the pilot speci�es a landing location. If nosuch command is received within a predetermined amount of time, the vehicle will revert intoautonomous mode and land at an automatically chosen location.

Depending on how far in advance the lidar scan can be conducted and transmitted, teleopera-tion may result in penalties in propellant consumption. As a worst-case scenario, consider thesituation of a 2 megapixel, 16 bit uncompressed lidar scan which is not conducted until directlyoverhead of the landing target. The vehicle must then wait in hover throughout the full timefor relay of information and pilot decision. Via the omnidirectional communications systemon S-band, it would take approximately 3.6 minutes to transmit the complete scan, consumingfar more propellant than could likely be allotted. If the scan were instead transmitted viathe high gain system at 100 mbps, however, transmission time would be approximately 0.3seconds, resulting in only 0.52 m/s of Δv consumption. Additional delays may occur due toantenna pointing requirements. If the remote pilot is located on Earth, there will be a furtherdelay of approximately 2.5 seconds (Δv = 4 m/s) due to the round trip light time. The �nalmajor source of delay would be the decision time taken by the human to inspect the lidar scanand designate a landing site. If the human decides quickly (within 5 seconds), approximately 8m/s of Δv will be consumed. Thus, teleoperation can be expected to carry a total Δv penaltyno larger than several tens of meters per second.

134

9.6 Navigation (Nate Niles)

Hard navigation requirements for the vehicle were not given. However, it was decided that a�nal position estimation uncertainty of 100 meters (3 sigma) at the end of each �ight wouldbe desirable. This should allow landing sites to be planned fairly close to the locations of theactual mission sites, giving the crew minimal transit times between the two and easing thetransportation of equipment and samples.

Given that the vehicle requires access to one of the LRS's prior to the mission, and that the�ights are on the order of 3-5 minutes, access to an LRS should be available upon landing. Thevehicle's position can then be determined within the limits of the LRS's ranging capabilityupon arrival at the landing site.

9.6.1 Position Estimation

There are very few ways to accurately determine one's position relative to an unseen reference(e.g., the origin of the Earth-Centered Inertial frame) quickly. This is the reason that GPS hasbeen implemented for Earth-based navigation, and also the reason that ranging is a plannedfeature for the LRS's. The Alshain vehicle will make use of the LRS ranging capability todetermine its position on the lunar surface, and inertial measurements to track its trajectoryduring �ight.

There will be a command interface to allow the vehicle estimated position to be set to aspeci�ed value.

Initial Fix

The achievable accuracy of position estimation using LRS ranging combined with �landmarktracking� is currently given as approximately 10 meters (NASA, 2007). This process is statedas requiring several minutes to settle. At the base, this position can be corrected by knownaltitude and ranging to the LCT's.

Position estimation with LRS ranging and �landmark tracking� involves a process by whichthe inertial position given by LRS ranging is used in conjunction with measurements to localterrain features (camera and/or lidar) to resolve the vehicle's position on a terrain map.

In-Flight

Since the nature of future lunar terrain data is unknown at this time, it would be di�cult todesign a system that relies heavily on it. Therefore, the navigation system has been designedto operate on only inertial feedback during �ight. For this reason, high accuracy IMU's arerequired to achieve the desired 100 meter landing position uncertainty (see 'Reference Models'section).

It is possible that a camera, lidar or radar system could be used in conjunction with terraindata to provide position and/or velocity updates to a �lter for the IMU data in-�ight, but thedesign of such a system would be highly dependent on the speci�cs of the terrain data andthe ability to store and process it in real-time.

Final �x

Once the vehicle has landed, its position will be determined using the same process as usedinitially, based on the availability of the external services (LRS and LCT). The worst case isagain given as approximately 10 meters (3 sigma) using LRS ranging and terrain matching.This is done to facilitate the most e�cient and accurate accomplishment of the mission.

135

9.6.2 Attitude Estimation

The Alshain vehicle has no speci�c requirements for attitude accuracy, other than those derivedfrom position accuracy. In order for inertial measurements to accurately re�ect the vehicle'schange in position, it is imperative that the attitude be known initially and that changes inattitude are included in the trajectory propagation.

Sources of attitude knowledge include surface features and celestial bodies. Knowledge ofsurface terrain alone, however, will probably not be su�cient for attitude estimation (althoughit could be used with a camera system for attitude rate estimation). That leaves celestialbodies, including the Earth, sun and stars as three distinct data sources.

The Earth will not necessarily be in view for all parts of every mission for which Alshain hasbeen designed. It is also subject to phasing in the same way that the moon is as seen fromEarth, complicating measurements.

The sun will also not necessarily be in view at all times, and even if missions were designedaround sun visibility, a sun sensor only provides a vector measurement (pitch and yaw posi-tions), and it must have line-of-sight to the sun to give a useful measurement.

This leaves the stars as reference objects and a star camera or �star tracker� as a device withwhich to determine the vehicle's attitude. By calculating vector positions to several objectsin their �eld of view, and determining their relative orientations, a star tracker can give roll,pitch and yaw orientations (although roll orientation accuracy is typically much less than thatof pitch and yaw). Star trackers make use of a catalog of celestial objects that is large enoughto be e�ective regardless of which objects are in their �eld of view, although they will beblinded by the sun and possibly the Earth. For this reason, two star trackers will be placedon the Alshain vehicle, pointed in di�erent directions (although both generally pointed up).Additionally, if neither star tracker is blinded or obstructed, then the vector measurementsfrom each unit can be combined to produce a higher-accuracy estimate than either unit alonecould produce (Markley, 2001).

To reduce the possibility of errors due to lunar dust buildup on the optics, covers will be putover the ba�es when the star trackers are not in use.

There will be a command interface to allow the vehicle estimated attitude to be set to aspeci�ed value.

Initial Fix

Attitude determination using the star trackers will begin several minutes before �ight. Ifneither star tracker has a clear view of the star �eld, a noti�cation will be sent to the crew andthrough telemetry so that the attitude may be set via the command interface, or the vehiclemay be moved manually to a better location.

In-Flight

Vehicle attitude will be determined in-�ight using inertial propagation only. This will allowthe star trackers to be covered for all but a few minutes of every mission, reducing the build-upof lunar dust.

136

Final Fix

Given the short duration of the �ights, and the low angular random walk rates of high-accuracyIMU's, the �nal propagated attitude will be used as the �nal �x after each �ight. (Note thata new �initial �x� will be taken before the return �ight of an out-and-back mission)

9.6.3 Flight Path Propagation

Since there will nominally be four IMU's in operation, the mean values of the four measure-ments will be used to propagate the vehicle position and attitude. This should serve to reducethe impact of their individual errors. This average could be weighted once the characteristicsof each individual IMU are known (after manufacturing and test).

9.6.4 Components

The navigation system will consist of the following components:

Table 51: Navigation Components

Multiples of each unit are required for redundancy, with the added bene�t of improved accu-racy.

Layout

The IMU's will be mounted in the main avionics bay. They will be mounted such that theiralignment with the body frame is constant, regardless of vibrations or temperature �uctua-tions.

The star trackers will be mounted on the upper surfaces of the vehicle such that their viewis unobstructed by permanent vehicle structures, and their alignment with the vehicle bodyframe is constant regardless of vibrations or temperature �uctuations. They will requireautomatic covers, so that the crew does not need to access them prior to each �ight.

The LRS Ranging Receiver antennas will be mounted such that they have an unobstructedview of the star �eld, and otherwise as necessary for their operation. The associated electronicswill be mounted in the avionics bay. The details of such a receiver, or if dedicated hardwareis even required, is not known at this time.

Reference Models

Reference models were chosen for the IMU's and star trackers. The actual vendors and modelsare not as important as the combination of characteristics constituting a reference that will beused to analytically demonstrate su�cient performance of the navigation system. Performancecharacteristics for the various components are given below in Table 1.

137

The reference IMU is the Honeywell HG9900 IMU. It incorporates digital laser gyros andaccelerometers24.

The reference star tracker is the Galileo Avionica A-STR Autonomous Star Tracker25.

9.6.5 Error budget

The most conservative method of error budgeting is to simply add the impact of each uncer-tainty and take the sum as the total. However, this tends to overestimate the total, becausethe uncertainties given for devices are vector magnitudes whose direction is unknown. Takingthe sum as the total assumes the worst case � that all of the errors are aligned in exactlythe same direction. In reality, we expect them to be completely independent, so a much lessconservative approach would be to take the root sum square of all of the values as the to-tal. A slightly less optimistic approach is to categorize the uncertainties by the frequency oftheir variation, take the root sum square within each category and then take the sum of thecategories.

The error budget for the Alshain navigation system uses the last of the three approachesabove. Star tracker and LRS ranging errors were taken as low frequency variation becausethese measurements are only taken once per �ight, so they are e�ectively constant in thisapplication. Gyro angular random walk (noise) is the only high frequency uncertainty givenfor the reference models used, so it was categorized separately.

Table 52: Navigation system uncertainties

24http://www51.honeywell.com/aero/common/documents/myaerospacecatalog-documents/MilitaryAC/HG9900_IMU.pdf25http://www.selex-sas.com/EN/Common/�les/Galileo_Avionica/Relazioni_Esterne/Scheda_Prodotto_2/Space_2/A_STR.pdf

138

Calculations

Initial Conditions

Initial position uncertainty is carried through the propagation unchanged.

Initial attitude uncertainty results in traveling in the wrong direction. For a given distancetraveled r at an angle θ the resulting position error δ is:

Accelerometer Bias Error

Accelerometer bias error results in an unknown, but constant, acceleration being applied tothe vehicle at all times. For a given time traveled t with unknown acceleration abias, theresulting position error δ is:

Accelerometer Scale Factor Error

Accelerometer scale factor error results in an unknown acceleration being applied to the ve-hicle that scales linearly with the measured acceleration (zero acceleration being the state offreefall). For a given time traveled t at an average acceleration (magnitude) a with an unknownacceleration scale factor asf, the resulting position error δ is:

Gyro Bias Error

Gyro bias error results in an unknown, but constant, angular velocity being applied to thevehicle at all times. For a given average velocity (magnitude) with unknown angular velocityωb over a time t, the resulting position error δ is:

Gyro Scale Factor Error

Gyro scale factor error results in an unknown angular velocity being applied to the vehicle thatscales linearly with the measured angular velocity. For a given average velocity (magnitude)with unknown angular velocity scale factor ωsf at an average angular velocity (magnitude)over a time t, the resulting position error δ is:

Gyro Angular Random Walk

Angular random walk is an error associated with random noise on a measurement. Even ifthe noise is a zero mean process, the sum at any particular instant is most likely not zero,and so angular random walk is modeled as an unknown angular velocity being applied to thevehicle that varies with the square root of time. For a given average velocity (magnitude) withunknown angular velocity ωrw over a time t, the resulting position error δ is:

139

Results

Table 53 below gives examples of the error budget propagated for two sample missions thathave been considered. One mission is a 10 km (one-way) trip to the center of Shackleton crater,and the other is the maximum range mission that Alshain has been designed to accomplish �a 57 km (one-way) trip to the center of Shoemaker crater. The values (t, a , v̄ andω̄ ) used inthe calculations were taken from an iterative simulation that incorporated acceleration times.

140

Table 53: Sample Uncertainty Propagations

9.7 Vehicle Status Monitoring and Fault Handling (Nick D'Amore)

The vehicle shall monitor all critical parameters to enable identi�cation and handling of faults.The two primary vehicle systems requiring status monitoring are the power and propulsionsystems. Propellant levels will be monitored to allow for estimation of remaining Δv sothat a low-propellant warning may be issued if necessary. Mass of propellant remaining mayalso be a useful quantity for estimation of the vehicle's center of mass for control purposes.One technique for monitoring cryogenic propellant levels involves radio frequency gaging26.Furthermore, the propellant feed system will be monitored to enable diagnosis of faults inthat system. The pressure will be monitored before and after all valves and regulators toidentify component failures. The power system will require monitoring of battery and fuel cellvoltages, along with the temperatures and pressures of the oxygen and hydrogen entering thefuel cell system.

In addition to monitoring vehicle status, the vehicle will also relay crew status via the commu-nications system. These data will consist of crew heart rate and other physiological parametersas measured by instrumentation built into the space suit. This information will be transmittedfrom the vehicle to mission control either directly or via relay satellite.

9.7.1 Fault Tolerance

The vehicle will have four �ight computers executing all instructions in parallel. In the eventof a discrepancy between the four computers, a voting architecture similar to that used onthe Space Shuttle will be employed to identify the faulty computer. In the event of a single

26http://academy.grc.nasa.gov/2008/research-projects/rf-mass-gauging-on-propellants-in-low-gravity

141

computer failure, the remaining three computers will still concur and the faulty fourth com-puter will be removed from active status. Should a second fault occur, two computers will stillconcur and the faulty computer will still be identi�able. In the unlikely event of a two-twosplit (i.e., two computers fail in precisely the same way, producing identical incorrect results)the correctly functioning pair of computers would not be readily identi�able. The system maynot be able to tolerate two faults under this unlikely set of circumstances.

Fault tolerance in the �ight navigation system is similarly achieved through the use of fourIMU's. The mean values of the four measurements will be used as the measurement by whichthe vehicle position and attitude are propagated. A threshold value will be set for the varianceof the individual measurements, and exceeding this limit will indicate failure of a unit and causeits future measurements to be rejected.

Redundancy in the command and data handling system is achieved through redundancyof components. Three multiplexer/demultiplexer (MUX/DEMUX) units will be employed;and all critical instruments will be connected via multiple paths to ensure that up to twoMUX/DEMUX failures will not result in inability to use these instruments.

While three or four-string systems are employed for critical avionics components such asthe �ight computers and IMU's, in many cases robustness to component failures is achievedthrough alternate systems. In the event of failure of the sole lidar unit, for example, pilotjudgment will be required to visually select a landing site free of hazards (Alternatively, anemergency return to the pre-prepared base landing site is possible). Two star trackers arecarried to ensure that an attitude �x can be acquired prior to launch in the event of a failure.Thus, the mission can continue in the event of loss of one star tracker. Star trackers are notrequired during �ight, so loss of a second star tracker, even during �ight, would not endangerthe crew and would at most be grounds for a mission abort. Similarly, it is shown in a previoussection that the vehicle can reach its destination with adequate precision using only inertialnavigation during �ight. Thus, loss of navigational references during �ight will not result inloss of mission as long as the inertial navigation system remains functional. Moreover, viaranging to the LCT, a return to base will be possible even in the event of a signi�cant loss ofnavigational functionality.

Because the communications system is critical to crew safety, it too was designed to be two-fault tolerant. Two high gain antennas are carried as the primary means of direct-to-earth(DTE) communications while on the ground. In the event that both dishes fail, the omnidi-rectional communications system is su�cient for low data rate voice communications DTE.During �ight, the omnidirectional system will be the primary means of communications dueto the di�culty of properly pointing the high gain system while moving. Thus, at least threeomnidirectional antennas will be required to ensure that communications are available evenin the event of two faults.

9.8 Control (Nate Niles)

As shown in Figure 107, the control system will accept guidance in the form of velocity, accel-eration, attitude and attitude rate. The output will be a thrust vector and a thrust momentfor use in a thruster �ring algorithm. This algorithm will determine distribution, timing andthrottling across the array of available thrusters, as appropriate. Details on controlling thepropulsion system are unknown at this time, so the thruster �ring algorithm is left as futurework.

142

Figure 107: Control System Diagram

10 Power (Arber Masati)

The power system for the Alshain comprises of two fuel cells, a set of Li-ion Phosphate(LiFePO4) batteries, and a set of Lithium/Carbon Mono�ouride (CFx) batteries. One fuelcell can supply enough power to complete a mission and half of the phosphate batteries cansupply the power needed to �y. The Carbon Mono�ouride batteries are non-rechargeablebatteries for use only in emergencies. The fuel cells and batteries connect to a power manage-ment and distribution unit (PMAD) which controls the operation of the fuel cells, acts as adc-to-dc converter, and automatically switches to a working power source in the unlikely caseof a failure.

Figure 108: Power Overview

143

10.1 Power Requirements

The power system must provide the necessary power and energy at all times during a mission.Because of its importance to the success of the mission and the survivability of the astronautsthe system must be two faults tolerant with one fault being mission safe and two faults beingcrew safe. Furthermore, the mass of the system must be minimized in order to save moneyand fuel. To better analyze the power requirements, a mission was split in three categories:in-�ight, landed at mission site, and 24 hour contingency.

10.1.1 In-Flight Power Requirements

During the �ying phase of the mission, the Alshain draws more power and operates moreelectronic equipment than the other two phases. Table 54 shows the equipment that areoperating during �ight and their respective power draw.

Table 54: Power Requirements

Flying is a critical time in the mission, to assure safe arrival at the landing site all of thesesystems are operating for the duration of �ight. Propulsion valves are listed as pulse loadsbecause they operate once to open and then a second time to close. Therefore, the highestpower needed for the Alshain to �y is 1560 Watts, including a 30% margin.

10.1.2 Landed Power Requirements

After the Alshain lands at the speci�ed site, and while the astronauts are completing theirmission, some of the electronics turn o� and the vehicle remains in a high-power stand-bymode. Table 55 lists the operating equipment during the landed phase of the mission.

144

Table 55: Landed Power Requirements

The vehicle is not moving during the landed phase so one IMU and two computers are turnedo� along with LIDAR, radar, and video cameras. Star trackers will be used for 10 minutes atthe end of the mission right before takeo�. Because of their low power and energy draw, theyare in the chart for completeness but are not included in the total power calculations. Thetotal power needed during the landed phase of the mission, including a 30% margin, is 720Watts.

10.1.3 Contingency Power Requirements

In the unlikely event that Alshain su�ers an unrecoverable fault and cannot return to base,preparations have been made to allow for crew survival. During a contingency scenario,Alshain will be placed in a low-power stand-by mode, in which only crew critical systems willbe powered. Table 56 lists those systems and their respective power draws.

Table 56: Contingency Power Requirements

The number of active units is drastically cut down to those that are necessary for survival.Life support is the power needed to run the astronaut's suits. The astronaut suits have enough

145

power for 8 hours, so in order to survive for 24 hours their suits need an alternative powersource. The total power drawn during the 24 hour contingency is 600 Watts, including amargin.

10.1.4 Summary of Power Requirements

To summarize, the power drawn for each of the scenarios is 1560 Watts (in-�ight), 720 Watts(landed), and 600 Watts (contingency). The energy needed during those scenarios is calculatedby multiplying the power by the amount of time that power is drawn. Table 57 lists the powerand energy for the three scenarios.

Table 57: Summary of Power Requirements

The calculations are conservative and assume the longest possible �ight and mission time ofsix minutes and eight hours respectively. For a mission of this duration, the power systemmust provide a wealth of power during �ight and a wealth of energy while landed. On topof that, the power system must be able to provide the 14kWh needed for the astronauts tosurvive in case of an emergency. The combination of batteries and fuel cells allows the systemto meet the power, energy, and two fault tolerance requirements.

10.2 Li-ion Batteries

In recent years, Li-ion batteries have become competitive in space applications because of theirspeci�c power and energy, and small size. As already stated, the Alshain is using two typesof Li-ion batteries, rechargeable Li-ion phosphate (LiFePO4) batteries and non-rechargeableLithium/Carbon Mono�ouride (CFx) batteries.

10.2.1 Rechargeable Batteries

Lithium-ion phosphate batteries are rechargeable batteries that have excellent power densities.In a recent publication, �Battery materials for ultrafast charging and discharging� Kang andCeder show that a modi�ed phosphate battery can achieve 100 kW/kg and 25 kW/liter. Thebattery capacity depends on the rate of discharge and is plotted in �gure 109. Alshain hastwo sets of Li-ion phosphate batteries; one of them can provide enough power for the longestpossible �ight of 6 minutes. Therefore each battery operates at the shown 10C line and hasa capacitance of about 140 Ah/kg. Furthermore, the batteries can be charged and dischargedmultiple times with little a�ect to the total capacity as shown in �gure 109.

146

Figure 109: A rate of nC corresponds to a full discharge in 1/n h (Courtesy Kang, B. and Ceder, G. �Batterymaterials for ultrafast charging and discharging� Nature Vol. 458. March 12th 2009: pg. 190-193)

To summarize, rechargeable Li-ion phosphate batteries can charge and discharge fast withouta�ecting their total capacity. Alshain is equipped with two half-kilogram cells of Li-ion phos-phate batteries; each cell has nine batteries and operates at around 28 volts. These batteriespower the vehicle during �ight.

10.2.2 Non-Rechargeable Batteries

In case of an emergency where both fuel cells are inoperable and �ying back to base is im-possible, preparations are needed to allow for crew survival. From the power and energycalculations, the necessary energy is 14kWh. CFx batteries have a high speci�c energy andenergy density ratings as shown in table 58.

Given the power and energy requirements and the speci�cations of the CFx batteries, thebattery mass required for a 24 contingency is 24 kg. A 1.1 kg cell of eleven batteries in series

147

Table 58: Non-Rechargeable Batteries (Courtesy http://www.quallion.com/sub-tc-primary.asp)

operates at around 28 volts. A total of 22 cells supply the needed power for 24 hours. However,batteries cannot discharge all of their energy, taking into account an 80% depth of discharge;four more cells are added to Alshain bringing the total to 26 cells weighing 28 kg.

10.3 Fuel Cells

A fuel cell is an electrochemical device in which oxygen and hydrogen transform to water. Thetransformation process produces electricity and is essentially the reverse of electrolysis, wherewater transforms into oxygen and hydrogen when current runs through the water. Figure 110below is a basic picture of how a fuel cell works. Oxygen and hydrogen �ow into the fuelcell, electrons from the hydrogen go through the circuit and reunite later with the oxygen andhydrogen to form water.

Figure 110: Fuel Cell (Courtesy http://www.fuelcells.org/basics/how.html)

10.3.1 PEM Fuel Cells

Polymer Electrolyte Membrane, or, Proton Exchange Membrane, or simply PEM fuel cells,operate with gaseous hydrogen and oxygen at low pressures and at a temperature of 60 degreesCelsius. NASA has stated that they are using PEM fuel cells in project constellation because�This fuel cell chemistry is smaller and more e�cient than previous types such as the alkalinefuel cells used on the Space Shuttle. Its compact size helps reduce the overall mass and volumeof the spacecraft's power system.�

148

Given that Alshain is part of the Constellation Program, the fuel cells powering the vehiclewill be of the PEM type. Table 59 lists some properties the fuel cells.

Table 59: PEM Fuel Cells

Although PEM fuel cells are NASA's choice for constellation, these fuel cells have never beenin space; however, they have been researched and used considerably on the ground, settingtheir technology readiness to level six.

10.3.2 Fuel Cell Sizing

Alshain consumes 700 Watts when landed at the mission site. The fault requirements statethat the mission must continue if there is only one fault; therefore, Alshain is equipped withtwo fuel cell with the ability to operate continually at 700 W and 50% e�ciency. Figure 111is a polarization curve showing the operating voltage and its dependence to amperes per cellarea for both air and pure oxygen supplies.

Air breathing fuel cells generate less power because oxygen makes up only 21% of the incominggas. Furthermore, a pure oxygen system does not have to deal with inert gasses such asnitrogen and carbon dioxide, making a pure oxygen system more e�cient. The e�ciency of afuel cell depends on the fuel cells operating voltage. The maximum possible voltage per cellis 1.482 volts however due to entropy and mechanical ine�ciencies fuel cells can never reachthat value. The e�ciency of the cell is its operating voltage divided by 1.482 volts.

482.1

V=η

Alshain's fuel cells will be operating at 50% e�ciency, so each cell will produce .741 volts. Theelectronic components will be operating at 28 volts; therefore, the fuel cell will produce 700Watts at 28 volts and 25 amps using 38 cells in one stack. From the polarization curve, thearea of each fuel cell is 35 cm2 (6 cm by 6 cm); each cell is .3 cm in width so the entire stackwill be 11.4 cm long. Including the casing, the fuel cell will have dimensions of 15 cm by 10cm by 10 cm. Using the weight per power estimation relation and adding a margin, each fuelcell has a mass of 3 kg.

149

Figure 111: Fuel Cell Sizing Graph (Courtesy Barber, Frano. PEM Fuel Cells Theory and Practice. ElsevierAcademic Press, 2005. pg 58)

10.3.3 Reactants Supply

For the fuel cells to operate oxygen and hydrogen must be supplied to them. The two fuelcells draw hydrogen and oxygen from the main line. Faraday's Law gives the amount of fuelused for a fuel cell; N is the molar mass rate, I is the current running through the cell, and Fis Faraday's constant, 96,485 Coulombs/electron-mole.

Table 60 shows the calculations for the mass �ow rate of the reactants and the product water.Each fuel cell will be producing 2.55 kg of water in an eight-hour period when operating at700 Watts.

150

Table 60: Reactants Supply

The reactants are drawn from the main lines; valves and pressure regulators ensure that thereactants are at the correct pressure of two atm. The main lines will be pressurized close tothe pressure of the main tanks; the pressure drop from 2 MPa, at the lines, to 2 atm, at theregulator, causes the reactants to �ash vaporize. The reactants are still too cold and cannot besent through the fuel cell; but they can be heated using the waste heat of the fuel cell. Eachfuel cell operates at 50% e�ciency and generates 700 Watts of waste heat. The heat removedby a �uid can be found using the �uid mass �ow rate, temperature di�erence, and constantpressure speci�c heat coe�cient.

∫ ⋅=2

1

.

*T

T

dTcpmQ

The speci�c heat coe�cient varies with temperature, but it can be approximated as a sec-ond order polynomial function, which is used to calculate the integral. Table 61 gives theapproximation values of cp for hydrogen and oxygen in Joules per mole-Kelvin.

Table 61: Reactants Supply Approximation

151

The mass �ow rate for oxygen and hydrogen is given in table 60. T1 is assumed to be theboiling temperature of each �uid (20 K for H2, 90 K for O2); T2 is the operating temperature ofthe fuel cell, 333 Kelvin (60*C). The amount of waste heat removed by heating the reactantsto 60*C is 21 Watts for oxygen and 44 Watts for hydrogen, which adds to 65 Watts. Theremaining 635 Watts is o�set to the thermal control system discussed in section [Add section,and numbers].

10.3.4 Water Management

A fuel cell will transform hydrogen and oxygen into water, which will form at the cathodeside. If too much water is present, the fuel cell drowns and stops working, until the water isremoved. To avoid drowning the fuel cells, oxygen will be supplied in excess in order to pushthe water out of fuel cell. Nominally, a fuel cell needs an 8:1 oxidizer to fuel ratio; however,a 10:1 ration is used to remove water. The water and excess oxygen, upon exiting the fuelcell, will be at a temperature and pressure of 60*C and two atm, and will �ow to the watertank. A fuel cell running for 8 hours will create 2.55 kg of water and use .57 kg excess oxygen.Including the excess oxygen, for 8 hours of operation, Alshain draws .29 kg of Hydrogen, 2.84kg of oxygen, and creates 2.55 kg of water.

10.4 Contingency Options

Safety is always the �rst priority. By NASA standards, if an accident were to occur and theastronauts were stranded away from base, the vehicle must provide them with enough powerfor 24 hours. Alshain is designed to meet and exceed that requirement. In this section, theunderlying assumption is that the thrusters are no longer working and the astronauts cannotwalk back to base.

In the event that fuel is available, one of the fuel cells can provide the necessary power. Theamount of fuel needed to power the necessary equipment for 24 hours is 9.4 kg in a 10 to 1ratio; 9.4 kg is 1% of the total fuel (940 kg). If, following a fault, 10% of the fuel (10 kg) wereavailable, a fuel cell could provide power to the crew for more than a week; at which point,food and water dictate crew survival.

In the event that fuel is unavailable, the fuel cells can no longer provide power to the astronauts.The only remaining option is to use the non-rechargeable CFx batteries. A 30% margin isalready built in the calculation for the batteries, along with the calculation for complete 24hour communication. The transceiver, antenna and WLAN will not be operating for the entire24 hours; however, the energy calculations have the communication systems on at all timesfor an added margin.

10.5 Power Management and Distribution

The four power sources, two fuel cells and two sets of batteries, connect to a power managementand distribution unit (PMAD). The PMAD controls the operation of the fuel cells by selectingwhich of the two operates and controlling its power output. In the event of a failure, the PMADautomatically switches to the other fuel cell or batteries. The power from the batteries andfuel cells goes through buck-boost choppers that regulate the voltage going to the electronics.The mass of the unit is estimated at 14 kg.

10.6 Wiring

The fuel cells and batteries supply power to the PMAD unit at 28 volts and varying amps, thePMAD unit then connects to the di�erent electronic components. The phosphate batteriessupply 1.5 kW of power at 55 amps. Each fuel cell can provide more power, albeit at lower

152

e�ciencies. At 40% e�ciency, a fuel cell can provide 1050 Watts of power at 47 amps and22.3 volts. The non-rechargeable batteries provide emergency power, at 28 volts and 21 amps.Lastly, the wiring from PMAD to the individual electronics provides at maximum two ampsto each component. Table 62 shows the mass calculations for copper wiring. Alshain has 80components that need power, for which the wiring is approximated at half a kilometer. Thetotal wire mass is 7 kg.

Table 62: Wiring

10.7 Summary of Power System

To summarize, the system comprises of two 50% e�cient, 700-Watt fuel cells; two half-kilogramLiFePO4 battery cells; 26 CFx battery cells weighing 28 kg; and a PMAD unit for control.Each fuel cell draws reactants at a 10 to 1 oxidizer to fuel ratio. An 8-hour mission requires.29 kg of hydrogen, 2.84 kg of oxygen, and creates 2.55 kg of water. The system mass issummarized in table 63.

Table 63: Power System Summary

153

11 Design Reference Mission (Andrew Wilson)

The Alshain vehicle is designed with the capability to provide astronauts the range and capa-bility to perform scienti�c missions in lunar crators. Detailed here is a reference mission forthe Alshain Lunar Flying Vehicle traveling from the proposed Altair landing site on the rimof Shackleton Crater to the Basin of Shackleton Crater.

11.1 Mission Site

Alshain's mission will include a trip to the Shackleton Crater Basin (89.9º S, 0.0º E) locatedclosest to the Altair lunar base site. With Shackleton's dimensions of 19 kilometers in diameterand a -2 kilometer altitude, Alshain will transport astronauts 20 kilometers roundtrip (10 kmto a mission site and 10 km back to base) with a total time of �ight of just 3.7 minutes (3minutes, 42 seconds). Shown in �gure 112 is a 3D image of Alshain's proposed �ight path forthis mission, using the same color coordination seen earlier for a nominal �ight plan.

Figure 112: Flight Path to Shackleton Crater Basin

Each ballistic hop, whether it is into or out of Shackleton or any other crater, has its ownspeci�c trajectory and set of detailed requirements. For the reference mission to Shackle-ton Crater Basin, �gure 113 details both the outbound and return trajectories for Alshain'sballistic hop �ights.

11.2 Payloads Carried

Alshain was designed with a payload capacity of 1.49x0.86x1 meters, providing an e�ectivevolume of 1.28 m3, to allow for the transportation of a third incapacitated astronaut in theevent of a contingency mission. With a compiled database of 103 potential payloads anda volume that allows the transportation of 99% of these payloads, there are hundreds ofdi�erent e�ective payload combinations. Alshain's ability to carry several di�erent payloadsat once is shown here in the mission to Shackleton Crater Basin where the payload bay holds

154

Figure 113: Ballistic Hop Trajectories for Mission to Shackleton Crater

the following payload combination: Rock Collection Tool Kit, Mass Spectrometer, VAPoR,MUM, and regolith drill.

The Rock Collection Tool Kit is an assortment of collection tools all carried together on aLarge Tool Carrier weighing approximately 18.1 kilograms with dimensions of 0.86x0.86x0.54meters. Included on the carrier are: sample bags, hammer, tongs, small or large adjustablescoop, extension handle and rake.

The regolith drill for this mission weighs 13.4 kilograms, with dimensions of 0.58x0.24x0.12meters and 430 Watts of self contained power. It can obtain a continuous soil column up tothree meters in length while providing holes for the placement of scienti�c instruments.

Three of the chosen payloads for this mission to Shackleton Crater were chosen based oncharacteristics that deemed them useful on experimental missions during exploration. Thespeci�c Mass Spectrometer shown in �gure 114 is the latest in e�orts to miniaturize andimprove spectrometry technology. It is a High-Performance Quadrupole Mass Spectrometerand the world's smallest at 2.3 kilograms and dimensions of 0.295x0.18x0.095 meters developedby the Jet Propulsion Laboratory at the California Institute of Technology. The device canidentify chemicals by their molecular weight and would ideally be used to help the astronautsdetect and identify external leaks in order to maintain exploration operations using Alshain.

Figure 114: Quadrupole Mass Spectrometer

155

VAPoR (Volatile Analysis by Pyrolysis of Regolith) is a pyrolysis-mass spectrometer instru-ment currently funded by the NASA Lunar Sortie Science Opportunities (LSSO) and devel-oped by the NASA Goddard Internal Research and Development (IRAD) programs. VAPoR,shown in �gure 115, is a single contained unit that can handle lunar regolith surface or sub-surface samples to test for volatile contents such as water and oxygen, a top priority in thesearch for in situ resources on the moon. With compact dimensions of 0.2x0.43x0.51 metersit weighs in at 17 kilograms.

Figure 115: VAPoR mass spectrometer

The last experimental payload on board the mission to Shackleton Crater is the MUM or MarsUnderground Mole. It is currently being developed through the Mars Instrument DevelopmentProgram (MIDP). The MUM, shown in �gure 116, is a subsurface penetration device within situ infrared re�ectance and raman spectroscopic sensing capabilities with the intention offurthering the search for life on Mars by accessing subsurface samples. This would provide anideal means of obtaining subsurface lunar regolith samples to be tested by VAPoR on site andincrease an astronauts' productivity during EVAs. The MUM payload is 3.5 kilograms with alength of 0.5 meters and a diameter of 0.04 meters.

Figure 116: Mars Underground Mole

All �ve payloads combined is a total of 54.3 kilograms of equipment being transported duringboth legs of the mission. This leaves a maximum margin of 115.7 kilograms to be used for thereturn of lunar regolith or rock samples to base, limited only by the available volume left inthe payload area, which is just more than 50 percent. These �gures seemed to be ideal basedon all of the Apollo exploration missions that collected and returned lunar samples. Apollo11 was the �rst mission to return with 22 kilograms of samples, Apollo 12 returned with 34kg, Apollo 14 with 42 kg, Apollo 15 with 77kg, Apollo 16 with 96 kg and �nally Apollo 17with 111kg in total. This particular mission to Shackleton Crater, through the use of Alshain,allows for maximum testing and collection of lunar samples from one single location and isideal for the continuous search for in situ resources on the moon.

156

11.3 Delta V requirements (Adam Kirk)

11.3.1 Assumptions/Calculations

In order to obtain the trajectories shown in the reference mission, an amount of change invelocity, ΔV, is required. This ΔV is directly related to the amount of fuel required for thetrajectories through the Tsiolkovsky rocket equation:

=∆initial

finale m

mVV ln

where Ve is the exhaust velocity of the rocket engine, minitial is the initial mass of the vehicle,and m�nal is the mass post-burn.

The ΔV calculations for the reference mission were based on equations developed for ballistichops and glides on �at, airless bodies. A �at, airless body is a good approximation because themoon has no atmosphere and the distance traveled is small compared to the moon's radius.Also, the small error that is introduced by the �at-moon approximation is conservative sincethe curvature would decrease the amount of ΔV necessary for a given distance. For ballistichops, it was assumed that two impulsive burns were made to launch and land the vehicle. Forglides, it was assumed that one horizontal impulsive burn was made with a constant verticalburn to balance out gravity for the duration of the �ight. Appendix A.2 includes a derivationof the ballistic hop and glide equations.

For the reference mission, the amount of ΔV was calculated by dividing the mission into threestages. The �rst stage was a ΔV for a ballistic hop 10 km horizontally and -2 km verticallyinto Shackleton Crater. The second stage consisted of a 3 km �at glide along the bottom ofthe crater. This accounts for any ΔV required in �nding a suitable landing area. Finally, thethird stage was the reverse of stage 1. It was a hop 10 km horizontally and 2 km verticallyback to the base site.

11.3.2 Results

The total ΔV required for the reference mission to Shackleton Crater was calculated to beabout 697 m/sec. This is the amount of ΔV obtained when adding together the ΔV require-ments for the three mission stages. The �rst and third ballistic hop stages each required 250m/sec and the second propulsive glide stage required 197 m/sec.

11.3.3 Reference to MATLAB scripts

A MATLAB script was used to run the ΔV calculations that have been described. It has beencommented and included in Appendix A.3. It uses the �at, airless body ballistic hop and glideequations derived in Appendix A.2.

157

12 Conclusions

12.1 Costing Analysis (Theo Talvac)

The Alshain Lunar Flying Vehicle was designed to operate as support to the Constellationprogram. The Constellation program is planned for a return to the moon by 2020 and there-fore, the Alshain LFV will be designed to be available for launch in 2020. In order to estimatethe expected production and development costs of the Alshain LFV, several assumptions weremade. The �rst assumption is that the Alshain LFV will be capable of being launched onthe initial �ight of the Ares V in 2020. The second assumption states that a learning curveof 85% was used when calculating recurring production costs. The third assumption is thatwill be two Alshain Lunar Flying Vehicles at all times in case of rescue operations on thelunar surface. All of these assumption were entered into the two NASA Cost Models used tocalculate the cost of the Alshain program.

12.1.1 NASA Cost Models

As stated, two NASA Cost Models were used in determining the production and developmentcosts of two Alshain lunar �ying vehicles. The �rst was the NASA Spacecraft/Vehicle LevelCost Model. The second was the NASA Advanced Missions Cost Model. Each model requiresinput from the user in order to calculate the costs. There are two important inputs in bothcost models. The �rst input is the mission type. Neither of these two models currently havea lunar �ying vehicle as a selection for the mission type. Therefore, the �nal production anddevelopment costs of the Alshain program are a rough estimate based on the output of eachNASA Cost Model. The second input is the dry weight. The dry weight is the mass of thevehicle without fuel, consumables, science and research packages, and astronauts. The AlshainLFV dry weight, along with other calculated masses, is shown in section 10.5.

12.1.2 In�ation

The two NASA Cost Models output the cost estimates in 2004 dollars. In order to improvethe cost estimates, an in�ation calculator was used. The in�ation calculator accounted forin�ation rates and output the cost estimates in 2008 dollars. The in�ation calculator is foundon the United States Bureau of Labor Statistics government website.

12.1.3 Spacecraft/Vehicle Level Cost Model (SVLC Model)

The Spacecraft/Vehicle Level Cost Model takes the following input from the user: programname, mission type, dry weight (pounds or kilograms), quantity, and the learning curve. Theinput entered for the Alshain LFV is displayed below:

The mission type is a manned spacecraft because the Alshain LFV requires astronauts tooperate the vehicle. The quantity and the learning curve used in the model are assumptionsmade for the Alshain LFV. The SVLC Model produced the following estimated productionand development costs:

As previously stated, the costs are in 2004 dollars. Using the in�ation calculator, each costis adjusted to account for the in�ation rates. The production and development costs in 2008dollars are shown in Table 66.

158

Table 64: Costing Input

Table 65: Costing Results

Table 66: SVLC Development Costs

Result (in millions FY 2008 US$) Development 1,104 Production 128 Total 1,232

159

Table 67: AMC Input Data

Table 68: 2004 Costs

12.1.4 Advanced Missions Cost Model (AMC Model)

The Advanced Missions Cost Model takes the following input from the user: quantity, dryweight (pounds), mission type, Initial Operating Capability year (initial launch year), blocknumber, and di�culty. The input entered for the Alshain LFV is displayed below:

The quantity remains as stated in the assumptions, two. The dry weight is the same weightused for the SVLC model. The mission type chosen for this model is a lunar rover because thechoice of lunar �ying vehicle is not available. The Alshain LFV is not a lunar rover, however alunar rover is the closest type of mission available for selection. To account for the discrepancyin the mission type, the level of programmatic and technical di�culty (di�culty input) wasset to high. The change in the di�culty input should result in a closer estimate of the totalcost of the Alshain LFV. The Initial Operating Capability year entered was 2020, the sameyear listed in the assumptions. The block number represents the level of design inheritance ofthe vehicle. The block number entered was one because a lunar �ying vehicle has never beentested on the lunar surface. The block number and di�culty inputs should adjust for the factthat a lunar rover was chosen as a mission type. The AMC Model produced the followingestimated total cost of the lunar �ying vehicle:

After using the in�ation calculator to account for the in�ation rates, the total cost in 2008dollars is shown in Table 69.

The AMC Model does not output the production and development costs like the SVLC Model.Using the learning curve, the production and development costs were derived for the AMCModel. First, the total cost for a quantity of 1, 2, and 3 lunar �ying vehicles was calculatedusing the AMC Model. Afterwards, the di�erence in total cost between each quantity was

160

Table 69: AMC Total Costs

Result (in millions FY 2008 US$) Total 1,282

Table 70: AMC Development Costs

Result (in millions FY 2008 US$) Development 745 Production 537 Total 1,282

found. This di�erence along with the assumed learning curve of 85% was used to �nd thedevelopment (non-recurring) cost and production (recurring) cost of the Alshain LFV. Thesecosts are shown in Table 70.

12.1.5 Preliminary Estimate

A preliminary estimate of development and production costs can be made from analyzing thetwo NASA Cost Models. The preliminary estimate along with costs from each model is shownin Table 71.

To further improve the estimate, NASA Cost Readiness Levels were used. The cost readinesslevels are shown in Table 72.

The cost readiness level of the Alshain LFV cost estimate is at 4. This means that the estimateis preliminary and within 45% of actual costs. Using the cost estimate from Table 4 and the45% margin of error, the bounds of the cost estimate can be derived. The upper bound is1,885 ($M) and the lower bound is 715 ($M).

12.2 Reliability Fault Tree (Ryan Lebois)

The following analysis examines the overall system reliability of the Alshain vehicle by creatinga system level fault tree. It is analyzed in two di�erent modes: loss of crew, and loss ofmission. The top level event of the fault tree in each system is the failure of the Alshainvehicle; however, one is a system level failure that leads to the loss of the mission, while the

Table 71: Preliminary Estimate

Millions FY2008 US$ SVLC Model AMC Model Estimate Development 1,104 745 ~900

Production 128 537 ~300 Total 1,232 1,282 >1,300

161

Table 72: Cost Readiness Levels

CRL Description Margin of error (+/-)

9 End of project cost 0

8 Cost fit for every firm engineering and very firm budget commitments

5%

7 Cost fit for firm engineering and firm budget commitments 15%

6 Cost fit for PDG engineering decisions and PDR budget use 25%

5 Cost fit for preliminary engineering decisions and preliminary budget decisions

35%

4 Cost fit for very preliminary engineering decisions and very preliminary budget decisions 45%

other is a system level failure that leads to the loss of the crew. It then breaks down bysubsystem, and incorporates lower level events that could also lead to total vehicle failure.On the most basic level, reliabilities are assigned to parts based on TRL level and componentresearch. This is an area of high uncertainty that greatly limits the accuracy of the analysis.

12.2.1 Parts List

The table below summarizes the individual parts that make up the vehicle and their respectivereliabilities:

As stated previously, at this stage in the development of the Alshain lunar �ying vehicle,component reliabilities are highly speculatory, and limit the overall accuracy of the reliabilityanalysis. That being said, a reliability study of this nature still provides a useful baseline forsystem reliability as well as an ideal method of subsystem analysis.

12.2.2 Fault Tree Structure

The following shows an outline of the main fault tree structure. A summary is shown sincethe actual fault tree is an extremely large and complicated web of subsystems, gates, andconditional faults:

12.2.3 Subsystems

Each of the subsystems shown directly below the top level event in the above summary, repre-sents a critical system that would lead to the loss of crew if it failed during �ight operations.In the case of a loss of mission analysis, even minor failures within these critical subsystemsthat don't cause the systems themselves to fail could lead to the loss of the mission.

This structuring allows for sub analysis of the critical systems individually in order to deter-mine week points at any level within the main system.

162

Table 73: Component Reliabilities

Parts Number Reliability Computer 4 0.9998 HUD 2 0.9999 Fuel Cell 3 0.999 IMU 3 0.9999 Joystick 2 0.9999 Keypad 2 0.9999 Pressurant 4 0.9999 LH2Tank 2 0.9999 LIDAR 1 0.9999 LOX Tank 2 0.9999 Main Thruster 1 0.999 Pres Regulator 11 0.999 RCS Thruster 20 0.999 Radar 1 0.9999 Valve 12 0.999 Main Batteries 1 0.9999 Backup Batteries 1 0,9999

Figure 117: Fault tree structure

163

Table 74: Loss of Crew Weak Points

Table 75: Loss of Mission Weak Points

12.2.4 Monte Carlo Simulation

These fault trees were developed using the program Open Fault Tree Analysis. This soft-ware allows the ability to run fault trees through Monte Carlo simulations to determine theweaknesses of the system as well as the overall system reliabilities.

Early analysis was done by giving all components equal reliabilities in order to locate andanalyze single points of failure within the system. This type of analysis had two results:

� It showed the weakness of the propulsion feed system, which led to the restructuring of the pressureregulator and valve scheme as well as the cross feeding of both propellant and pressurant tanks.

� It showed the lack of redundancy in certain areas, which led to the addition of a fourth pressure vesseland additional RCS thrusters.

12.2.5 Results: Loss of Crew

Upon simulation of the fault tree built for loss of crew analysis, a total system reliability of99.6% is found. The weakest areas leading to this reliability are as follows:

NASA requires a 99.9% system reliability to be achieved prior to use, but this is only thereliability of a very preliminary design. However, we have exposed these areas of componentweakness within our analysis that could be used as focus points for future research e�orts.

12.2.6 Results: Loss of Mission

Upon simulation of the fault tree built for loss of mission analysis, a total system reliability of99.2% is found. In this case, the weakest areas leading to this reliability are as follows:

164

Table 76: TRL De�nitions

TRL Definition 1 Basic principles observed and reported: Transition from scientific research to applied

research. 2 Technology concept and/or application formulated: Applied research. Theory and

scientific principles are focused on specific application area to define the concept. 3 Analytical and experimental critical function and/or characteristic proof-of concept:

Proof of concept validation. Active Research and Development (R&D) is initiated with analytical and laboratory studies.

4 Component/subsystem validation in laboratory environment: Standalone prototyping implementation and test. Integration of technology elements.

5 System/subsystem/component validation in relevant environment: Thorough testing of prototyping in representative environment.

6 System/subsystem model or prototyping demonstration in a relevant end-to-end environment (ground or space): Prototyping implementations on full-scale realistic problems. Partially integrated with existing systems.

7 System prototyping demonstration in an operational environment (ground or space): System is at or near scale of the operational system.

8 Actual system completed and "mission qualified" through test and demonstration in an operational environment (ground or space): End of system development. Fully integrated with operational hardware and software systems.

9 Actual system "mission proven" through successful mission operations (ground or space): Fully integrated with operational hardware/software systems. Actual system has been thoroughly demonstrated and tested in its operational environment. �

12.3 Technology Readiness Levels (Theo Talvac)

Through the iterative process of designing the Alshain LFV, the technology readiness level ofevery component using on the vehicle was analyzed. For the analysis, the NASA TechnologyReadiness Level de�nitions were used. These de�nitions are listed in Table 76.

12.3.1 De�nition

12.3.2 TRL List (Theo Talvac, Mike Sotak)

Below is a summary of the TRLs compiled for the Alshain LFV:

12.4 Outreach (Ryan Lebois)

Throughout every year, NASA makes it a point to stress community outreach as a high priorityin all areas of the public. A widespread interest in space and space exploration is critical tothe success of programs such as NASA. Not only is it important to educate the public oncurrent missions and technology in order to gain approval, it is also important to interest andmarvel the youth, the generation which must one day continue the goals set forth today.

When government funding and investment come into play, so does public approval. A programthat loses all popularity with the public and becomes detached from the people, will not receivetheir interest when it comes to selecting world leaders or supporting policies necessary to thatprogram. For this reason, community outreach to the public at large is critical for NASA.

In addition, a focused outreach to the youth of the community is equally if not more vital toNASA's survival. Most students are not exposed to engineering at all through high school,and the science concepts involving astronomy and space exploration are again hardly covered.If NASA wants to maintain interest in the youth so that the college system in this countrycontinues to produce aerospace engineers and scientists that are well educated and ready to

165

Table 77: TRL List

166

take up the challenge of their �eld, then it needs to focus an e�ort on showing the marvels oftheir research and work to this young audience.

12.4.1 Overview

As part of the Alshain project, we made a concentrated e�ort to reach out to the youth of thecommunity, and engage them in an educational and interesting way. We had two goals for ouroutreach program including 100% class participation and 150+ hours of total outreach. Bothgoals were met or surpassed with a total of 198 hours of outreach from 15 events. At everystep along the way, we were able to discuss our work, the work of NASA, and why this shouldbe interesting to the people we engaged. Our e�orts consisted of:

� Sta�ng 3 of the Maryland Engineering Challenges in Baltimore, MD

� Judging 3 local science fairs in the Washington, DC area

� Presenting our work at 3 open houses for prospective engineering students of the University of Maryland

� Hosting an outreach presentation for the fraternity Zeta Psi

� Giving presentations of our work and demonstrations of our mock-up to thousands at Maryland Day

� Presenting our work to graduate students and members of industry during a Preliminary and CriticalDesign Review

12.4.2 Maryland Day

Our largest outreach function of the year was our demonstration at Maryland Day. MarylandDay is a celebration of the enormity of activities that go on at the University, and o�ers acrowd of 70,000+ current students, alumni, family and friends the chance to explore whatmakes the University of Maryland one of the leading research universities in the country. TheAlshain team was able to display our mockup for those interested in seeing demonstrations ofour loading and unloading techniques, as well as the scale of our vehicle. In addition, postersdisplaying work from throughout the semester were displayed for presentation with the mock-up. This way we were able to give our audience throughout the day a full understanding of theconcept and design of the Alshain as well as a hands on look at some of the hardware testing.

12.4.3 University Outreach

Our university focused outreach took on three fronts: Maryland Day (as already described),open house presentations to perspective engineering students, and a presentation to a campusfraternity. Each of these was focused on a di�erent audience, but all aiming to bring moreattention and interest to the engineering community.

Each spring the university holds three open houses where perspective students come and tourthe campus. Part way through the day they split into the schools that they are more interestedin. Our class took this opportunity to present brie�y on our project, as well as the class andaerospace engineering department in general to the visiting students. These open houseso�ered an engaging atmosphere where students and parents alike could ask us questions, andfeel comfortable that they were getting an honest response about our engineering program.

The fraternity event was more focused on showing the fun and interesting parts of engineeringto an audience that otherwise might not be exposed to these things. Our class's own fraternitymembers were able to get the event approved by their chapter as a university outreach eventfor the fraternity. By listening to our talk about Alshain, they were given the chance to explorethe engineering community and the work that we do.

167

Figure 118: Showing o� the Alshain Mock-up at Maryland Day

12.4.4 Community Outreach

Our outreach to the community focused on three separate areas; however, they were all relatedto the education of young students who might be interested in engineering. These three areaswere: sta�ng the Maryland Engineering Challenges, mentoring aerospace magnet students atParkland Middle School, and judging three science and technology fairs.

The three Maryland Engineering Challenges that we sta�ed are run each year in BaltimoreCounty, MD, and they encourage young students from local schools to engage in hands onengineering experience. The elementary school event was a Paper Airplane Challenge, themiddle school event was an Electric Cargo Plane Challenge, and the high school event was aHovercraft Challenge. We were able to sta� all three events, where we had a chance to talkto the students about their projects while at the same time presenting di�erent levels of ourown project. Reaching students at such a young age, and catching their attention with the funand engaging aspects of engineering does a lot to advance the �eld and increase the supportthat we receive.

Another opportunity that we took advantage of was to mentor students at Parkland AerospaceMagnet Middle School. Their classes were preparing their science fair projects this spring andrequested some students from the university to come and help them review and improve theirprojects. We were able to ask them thought provoking questions about the research and teststhat they had run in an attempt to help them improve their evaluation and analysis skills,as well as develop more meaningful experiments. We were informed later that these studentsswept the science fair winning every prize, and were very ready to attribute our mentors withmuch of the credit for their success.

168

Figure 119: Listening to students' reports at one of the Maryland Engineering Challenges

Figure 120: Measuring range of �ight at on of the Maryland Engineering Challenges

169

The science fairs that we judged in DC, Howard County, and Prince George's County allowedus to see some of the very outgoing students in the surrounding community. While judgingprojects, we were able to learn about the interests of the students as well as to steer thoseinterests towards applications to engineering. We awarded several prizes on behalf of theCapitol PC Users Group, and were able to talk to hundreds of students from schools throughoutthe DC area.

12.4.5 Additional Outreach

The last piece to our outreach program involved the full technical presentation of our project tograduate students and industry professionals at both Preliminary and Critical Design Reviews.While outreach to areas of the community that are not often exposed to engineering applica-tions is important, it is also important to disseminate information throughout the aerospacecommunity itself.

12.4.6 Outreach Summary

The following table shows a summary of our outreach activities, participation, and hours. Amore detailed outline of individual participation and hours is provided in Appendix 6.

Table 78: Outreach Summary

Event Student Participation Hours MD Engineering Challenges 8 48 Science Fairs 4 16 Parkland Mentoring 5 18 Open Houses 13 17 Fraternity Presentation 2 2 Maryland Day 25 89 Design Reviews 28 8

Totals: 100% 198

12.5 Mass Budget (Neal Vasilak)

12.5.1 Total Inert Mass Breakdown

170

Table 79: Total Inert Mass Background

12.5.2 Percentage Breakdown

�� !�� "#$�� "%&� �� #'�()��(�*��+� �!� �� #��+��

171

�� !��"�#�� $�� %� �� %� ��!��&��&�� '�(�� %� �� "&��)�� %� �� %� ��

172

A Appendix

A.1 Trade Studies

A.1.1 One vs Two Person Vehicle (Neal Vasilak)

The goal of this trade study was to determine whether a one person vehicle was more feasiblethan a two person vehicle.

Advantages of a One Person Vehicle

� 250 kg less inert mass

� At least 2 cubic meters less volume

Disadvantages of a One Person Vehicle

� Must send two vehicles per mission (NASA STD)

� Avionics doesn't scale to smaller vehicle

� Higher probability of failure with two vehicles

� Problems landing two vehicles in close proximity

Fuel Consumption

2 Person Vehicle 1 Person Vehicle Inert Mass: 1050 kg 800 kg each Fuel necessary for 114 km round trip: 764 kg 803 kg (2 vehicles) �

Saves 5% more fuel for a two person vehicle per mission than two one man vehicles.

Conclusions

Due to the 50 % less chance of failure with a two person vehicle over two one person vehicles,and a savings of 5% more fuel per mission, the two person vehicle was the chosen design.

A.1.2 Battery Mass Trade Study

This trade study focuses on the minimization of the power system mass. More speci�cally, thecomparison is that of di�erent Li-ion battery types and fuel cells. Di�erent Lithium batterieshave di�erent speci�cations, which table [#?] lists. Given these speci�cations, the mass of abattery-powered system is shown.

173

Though Lithium/Carbon Mono�ouride batteries o�er a low system mass, they are not recharge-able and therefore would need replacing after every mission. The most mass e�cient systemusing rechargeable batteries weighs 40 kg. However, batteries cannot discharge fully, and whentwo-fault tolerance is taken into account the mass of the battery system reaches 48 kg.

Fuel cells can provide the necessary power as long as there is fuel. Two fuel cells provideenough power to �y, while a third one provides fault tolerance. A system of three fuel cellshas a mass of 9 kg for the fuel cells, 6 kg of piping, and 4 kg of fuel. Thought the systemweighs 20 kg, it does not allow for production of power when fuel is not available. Shouldthe astronauts become stranded with no fuel, the 24-hour contingency requirement cannot bemet.

The current battery-fuel-cell combination system meets the 24-hour requirement by includingCFx batteries, and minimizes mass by using two fuel cells and just one kilogram of batteries.

174

A.1.3 Con�guration Analysis

Item Weight X-Pos Y-Pos Z-Pos Avionics 115 0 0 0.15

Fuel Cells 32.5 0 0 0.55 Dish 0 0 0 0 Dish 0 0 0 0

Load/Unload Device 0 0 0 0 Crew 1 --- -1 0 1.6 Crew 2 --- 1 0 1.6

Crew Shielding 0 0 0 0 Star Tracker 0 0 0 0

Rocket Nozzle 1 100 0 0 -0.25 Rocket Nozzle 2 0 0 0 -0.25 Rocket Nozzle 3 0 0 0 -0.25 Rocket Nozzle 4 0 0 0 -0.25

Payload to landing site --- 0 0 1.3 Return Payload --- 0 0 1.3

Empty Pressure Tank 40 1.1 1.1 0.25 Empty Pressure Tank 40 -1.1 -1.1 0.25 Empty Pressure Tank 40 1.1 -1.1 0.25 Empty Pressure Tank 40 -1.1 1.1 0.25

Empty LOX Tank 12 0 0.95 0.25 Empty LOX Tank 12 0 -0.95 0.25 Empty LH2 Tank 17 1 0 0.25 Empty LH2 Tank 17 -1 0 0.25 Compressed Gas 18 - - -

LOX Fuel 617 - - - LH2 Fuel 103 - - -

Full Pressure Tank 44 1.1 1.1 0.25 Full Pressure Tank 44 -1.1 -1.1 0.25 Full Pressure Tank 44 1.1 -1.1 0.25 Full Pressure Tank 44 -1.1 1.1 0.25

Full LOX Tank 341.5 0 0.95 0.25 Full LOX Tank 341.5 0 -0.95 0.25 Full LH2 Tank 57.5 1 0 0.25 Full LH2 Tank 57.5 -1 0 0.25

175

X-CG of vehicle X-CG from rocket center= 0

Pre-flight Post-flight Pre-flight Absolute

Value Case 1 0 0 0 0 Case 2 -0.046977764 -0.0892326 -0.046977764 0.046977764 Case 3 0.046977764 0.0892326 0.046977764 0.046977764 Case 4 0 0 0 0 Case 5 0 0 0 0 Case 6 -0.051849291 -0.108616944 -0.051849291 0.051849291 Case 7 0.051849291 0.108616944 0.051849291 0.051849291 Case 8 0 0 0 0 Case 9 -0.098586921 -0.195950359 -0.098586921 0.098586921 Case 10 -0.051849291 -0.108616944 -0.051849291 0.051849291 Case 11 0.098586921 0.195950359 0.098586921 0.098586921 Case 12 0.051849291 0.108616944 0.051849291 0.051849291 Case 13 -0.109369304 -0.243704305 -0.109369304 0.109369304 Case 14 -0.057848052 -0.138760407 -0.057848052 0.057848052 Case 15 0.109369304 0.243704305 0.109369304 0.109369304 Case 16 0.057848052 0.138760407 0.057848052 0.057848052 Case 17 0 0 0 0 Case 18 0 0 0 0

Post-flight Absolute Value

0 0 0 0 -0.0892326 0.0892326 0.046978 0.089233 0.0892326 0.0892326 0.046978 0.089233

0 0 0 0 0 0 0 0

-0.108616944 0.108616944 0.051849 0.108617 0.108616944 0.108616944

0 0 -0.195950359 0.195950359 -0.108616944 0.108616944 0.195950359 0.195950359

0.108616944 0.108616944

MAX CG Offset (X-dir)

(m) = 0.243704305 -0.243704305 0.243704305 -0.138760407 0.138760407 0.243704305 0.243704305 0.138760407 0.138760407

0 0 0 0

176

Item Weight X-Pos Y-Pos Z-Pos Avionics 115 0 0 1.05

Fuel Cells 32.5 0 0 1.6 Dish 0 0 0 0 Dish 0 0 0 0

Load/Unload Device 0 0 0 0 Crew 1 --- -1 0 1.6 Crew 2 --- 1 0 1.6




Empty Pressure Tank 40 0.36 0.36 0.5 Empty Pressure Tank 40 0.36 -0.36 0.5 Empty Pressure Tank 40 -0.36 0.36 0.5 Empty Pressure Tank 40 -0.36 -0.36 0.5

Empty LOX Tank 12 1.1 -1.1 0.55 Empty LOX Tank 12 -1.1 1.1 0.55 Empty LH2 Tank 17 -1.1 -1.1 0.55 Empty LH2 Tank 17 1.1 1.1 0.55 Compressed Gas - - - -

LOX Fuel - - - - LH2 Fuel - - - -

Full Pressure Tank 44 0.36 0.36 0.5 Full Pressure Tank 44 0.36 -0.36 0.5 Full Pressure Tank 44 -0.36 0.36 0.5 Full Pressure Tank 44 -0.36 -0.36 0.5

Full LOX Tank 351.4285714 1.1 -1.1 0.55 Full LOX Tank 351.4285714 -1.1 1.1 0.55 Full LH2 Tank 58.57142857 -1.1 -1.1 0.55 Full LH2 Tank 58.57142857 1.1 1.1 0.55

177

X-CG of vehicle

X-CG from rocket center= 0


Value Case 1 0 0 0 0 Case 2 -0.046339203 -0.0892326 -0.046339203 0.046339203 Case 3 0.046339203 0.0892326 0.046339203 0.046339203 Case 4 0 0 0 0 Case 5 0 0 0 0 Case 6 -0.051072523 -0.108616944 -0.051072523 0.051072523 Case 7 0.051072523 0.108616944 0.051072523 0.051072523 Case 8 0 0 0 0 Case 9 -0.09718173 -0.195950359 -0.09718173 0.09718173 Case 10 -0.051072523 -0.108616944 -0.051072523 0.051072523 Case 11 0.09718173 0.195950359 0.09718173 0.09718173 Case 12 0.051072523 0.108616944 0.051072523 0.051072523 Case 13 -0.107642626 -0.243704305 -0.107642626 0.107642626 Case 14 -0.056882821 -0.138760407 -0.056882821 0.056882821 Case 15 0.107642626 0.243704305 0.107642626 0.107642626 Case 16 0.056882821 0.138760407 0.056882821 0.056882821 Case 17 0 0 0 0 Case 18 0 0 0 0


0 0 0 0 -0.0892326 0.0892326 0.046339 0.089233 0.0892326 0.0892326 0.046339 0.089233

0 0 0 0 0 0 0 0

-0.108616944 0.108616944 0.051073 0.108617 0.108616944 0.108616944 0.051073 0.108617

0 0 0 0 -0.195950359 0.195950359 0.097182 0.19595 -0.108616944 0.108616944 0.051073 0.108617 0.195950359 0.195950359 0.097182 0.19595 0.108616944 0.108616944 0.051073 0.108617

-0.243704305 0.243704305 0.107643 0.243704 -0.138760407 0.138760407 0.056883 0.13876 0.243704305 0.243704305 0.107643 0.243704 0.138760407 0.138760407 0.056883 0.13876

0 0 0 0 0 0 0 0

MAX CG Offset (X-dir) (m) = 0.243704305 MAX CG Offset (Y-dir) (m) = 0

178

Item Weight X-Pos Y-Pos Z-Pos Avionics 100 -0.65 0 0.15

Fuel Cells 32.5 -0.65 0 0.4 Dish 9 -0.684 0.405 0.8 Dish 5 -0.685 -0.321 0.8

Load/Unload Device 0 0 0 0 Crew 1 --- 1.485 -0.485 0.727 Crew 2 --- 1.485 0.485 0.727



Payload to landing site --- 0.3 0 0.3 Return Payload --- 0.3 0 0.3

Empty Pressure Tank 48 0 -1.05 1.255 Empty Pressure Tank 48 0 1.05 1.255 Empty Pressure Tank 48 -1 0 0.35 Empty Pressure Tank 0 0 0 0

Empty LOX Tank 12 0.55 -1.05 0.55 Empty LOX Tank 12 -0.55 1.05 0.55 Empty LH2 Tank 17 -0.55 -1.05 0.55 Empty LH2 Tank 17 -0.55 1.05 0.55 Compressed Gas 0 - - -

LOX Fuel 0 - - - LH2 Fuel 0 - - -

Full Pressure Tank 54 0 -1.05 1.255 Full Pressure Tank 54 0 1.05 1.255 Full Pressure Tank 54 -1 0 0.35 Full Pressure Tank 0 0 0 0

Full LOX Tank 351.4285714 0.55 -1.05 0.55 Full LOX Tank 351.4285714 -0.55 1.05 0.55 Full LH2 Tank 58.57142857 0.55 -1.05 0.55 Full LH2 Tank 58.57142857 -0.55 1.05 0.55

179

X-CG of vehicle

X-CG from rocket center= 0


Value Case 1 0.203034853 0.365157485 0.203034853 0.203034853 Case 2 0.143073901 0.263168185 0.143073901 0.143073901 Case 3 0.143073901 0.263168185 0.143073901 0.143073901 Case 4 0.077228656 0.140740147 0.077228656 0.077228656 Case 5 0.193519136 0.378215097 0.193519136 0.193519136 Case 6 0.126879257 0.254965108 0.126879257 0.126879257 Case 7 0.126879257 0.254965108 0.126879257 0.126879257 Case 8 0.05298803 0.100825397 0.05298803 0.05298803 Case 9 0.077228656 0.140740147 0.077228656 0.077228656

Case 10 0.004588235 -

0.008954714 0.004588235 0.004588235 Case 11 0.077228656 0.140740147 0.077228656 0.077228656

Case 12 0.004588235 -

0.008954714 0.004588235 0.004588235 Case 13 0.05298803 0.100825397 0.05298803 0.05298803

Case 14 -

0.029406214 -0.09748042 -0.029406214 0.029406214 Case 15 0.05298803 0.100825397 0.05298803 0.05298803

Case 16 -

0.029406214 -0.09748042 -0.029406214 0.029406214

Case 17 -

0.075956474 -

0.196167084 -0.075956474 0.075956474

Case 18 -

0.121860806 -

0.362109253 -0.121860806 0.121860806

Post-flight Absolute Value 0.365157485 0.365157485 0.203038 0.365165 0.263168185 0.263168185 0.144667 0.266451 0.263168185 0.263168185 0.145066 0.267271 0.140740147 0.140740147 0.07724 0.140767 0.378215097 0.378215097 0.193524 0.378225 0.254965108 0.254965108 0.12906 0.260012 0.254965108 0.254965108 0.129603 0.261267 0.100825397 0.100825397 0.053009 0.100883 0.140740147 0.140740147 0.090024 0.169506

-0.008954714 0.008954714 0.024064 0.05176 0.140740147 0.140740147 0.091425 0.172603

-0.008954714 0.008954714 0.026825 0.057737 0.100825397 0.100825397 0.073749 0.155319 -0.09748042 0.09748042 0.039479 0.117491 0.100825397 0.100825397 0.075837 0.160566 -0.09748042 0.09748042 0.041632 0.122013

-0.196167084 0.196167084 0.075971 0.196197 -0.362109253 0.362109253 0.121872 0.362138

180

MAX CG Offset (X-dir) (m) = 0.378215097

X-CG of vehicle Y-CG from rocket center= 0


Value Case 1 0.001215371 0.002270451 0.001215371 0.001215371 Case 2 -0.021412535 -0.041693989 -0.021412535 0.021412535 Case 3 0.023956969 0.046648452 0.023956969 0.023956969 Case 4 0.001334642 0.002725451 0.001334642 0.001334642 Case 5 0.001334642 0.002725451 0.001334642 0.001334642 Case 6 -0.023622291 -0.050979955 -0.023622291 0.023622291 Case 7 0.026429309 0.057037862 0.026429309 0.026429309 Case 8 0.001479869 0.003408521 0.001479869 0.001479869 Case 9 -0.04626104 -0.094468938 -0.04626104 0.04626104 Case 10 -0.023622291 -0.050979955 -0.023622291 0.023622291 Case 11 0.048930324 0.09991984 0.048930324 0.048930324 Case 12 0.026429309 0.057037862 0.026429309 0.026429309 Case 13 -0.051294886 -0.118145363 -0.051294886 0.051294886 Case 14 -0.026340621 -0.065587393 -0.026340621 0.026340621 Case 15 0.054254625 0.124962406 0.054254625 0.054254625 Case 16 0.029470656 0.073381089 0.029470656 0.029470656 Case 17 0.001479869 0.003408521 0.001479869 0.001479869 Case 18 0.001660562 0.004548495 0.001660562 0.001660562

MAX CG Offset (Y-dir) (m) = 0.124962406

Post-flight Absolute Value 0.002270451 0.002270451

-0.041693989 0.041693989 0.046648452 0.046648452 0.002725451 0.002725451 0.002725451 0.002725451

-0.050979955 0.050979955 0.057037862 0.057037862 0.003408521 0.003408521

-0.094468938 0.094468938 -0.050979955 0.050979955

0.09991984 0.09991984 0.057037862 0.057037862

-0.118145363 0.118145363 -0.065587393 0.065587393 0.124962406 0.124962406 0.073381089 0.073381089 0.003408521 0.003408521 0.004548495 0.004548495

181


Fuel Cells 32.5 -0.42 0 0.55 Dish 9 -0.42 0.2 1.2 Dish 5 -0.42 -0.2 1.2

Load/Unload Device 0 0 0 0 Crew 1 --- 0 0 1.347 Crew 2 --- 1 0 1


Rocket Nozzle 1 100 0 0 -0.25 Rocket Nozzle 2 0 0 0 0 Rocket Nozzle 3 0 0 0 0 Rocket Nozzle 4 0 0 0 0

Payload to landing site --- 0.6 0.6 0.7 Return Payload --- 0.6 0.6 0.7

Empty Pressure Tank 48 0 1.85 0.5 Empty Pressure Tank 48 0 -1.85 0.5 Empty Pressure Tank 48 -0.9 0 0.75 Empty Pressure Tank 0 0 0 0

Empty LOX Tank 12 0.55 1 0.54 Empty LOX Tank 12 -0.55 -1 0.54 Empty LH2 Tank 17 0.55 1 0.54 Empty LH2 Tank 17 -0.55 -1 0.54 Compressed Gas 34 - - -

LOX Fuel 617 - - - LH2 Fuel 103 - - -

Full Pressure Tank 54 0 1.85 0.5 Full Pressure Tank 54 0 -1.85 0.5 Full Pressure Tank 54 -0.9 0 0.75 Full Pressure Tank 0 0 0 0

Full LOX Tank 351.4285714 0.55 1 0.54 Full LOX Tank 351.4285714 -0.55 -1 0.54 Full LH2 Tank 58.57142857 0.55 1 0.54 Full LH2 Tank 58.57142857 -0.55 -1 0.54

182

X-CG of vehicle X-CG from rocket center= 0


Value Case 1 0.077372654 0.150550918 0.077372654 0.077372654 Case 2 0.034218896 0.073187614 0.034218896 0.034218896 Case 3 0.080991581 0.164262295 0.080991581 0.080991581 Case 4 0.035897939 0.080521042 0.035897939 0.035897939 Case 5 0.026084396 0.060480962 0.026084396 0.026084396 Case 6 -0.024169247 -0.044142539 -0.024169247 0.024169247 Case 7 0.027430341 0.067216036 0.027430341 0.027430341 Case 8 -0.025484222 -0.049674185 -0.025484222 0.025484222 Case 9 -0.013169774 -0.019679359 -0.013169774 0.013169774 Case 10 -0.013849329 -0.021870824 -0.013849329 0.013849329 Case 11 0.084965653 0.180721443 0.084965653 0.084965653 Case 12 0.037750258 0.089487751 0.037750258 0.037750258 Case 13 -0.079891186 -0.174987469 -0.079891186 0.079891186 Case 14 -0.084487917 -0.200057307 -0.084487917 0.084487917 Case 15 0.028922742 0.075639098 0.028922742 0.028922742 Case 16 -0.026950518 -0.056790831 -0.026950518 0.026950518 Case 17 -0.014602829 -0.024611529 -0.014602829 0.014602829 Case 18 -0.08964591 -0.233511706 -0.08964591 0.08964591


0.150550918 0.150550918 0.094408 0.181323 0.073187614 0.073187614 0.066162 0.13234 0.164262295 0.164262295 0.098824 0.197837 0.080521042 0.080521042 0.069409 0.145601 0.060480962 0.060480962 0.02609 0.06049

-0.044142539 0.044142539 0.024176 0.044159 0.067216036 0.067216036 0.027436 0.067227

-0.049674185 0.049674185 0.025491 0.049692 -0.019679359 0.019679359 0.060847 0.122895 -0.021870824 0.021870824 0.063987 0.136581 0.180721443 0.180721443 0.103673 0.217661 0.089487751 0.089487751 0.07299 0.161815

-0.174987469 0.174987469 0.079893 0.174993 -0.200057307 0.200057307 0.08449 0.200063 0.075639098 0.075639098 0.028929 0.075651

-0.056790831 0.056790831 0.026958 0.056811 -0.024611529 0.024611529 0.067468 0.153696 -0.233511706 0.233511706 0.089648 0.233519

183




Value Case 1 0.054095919 0.101057318 0.054095919 0.054095919 Case 2 0.05662613 0.110261081 0.05662613 0.05662613 Case 3 0.05662613 0.110261081 0.05662613 0.05662613 Case 4 0.059404645 0.121309285 0.059404645 0.059404645 Case 5 0.000523389 0.001068804 0.000523389 0.000523389 Case 6 0.000550396 0.001187825 0.000550396 0.000550396 Case 7 0.000550396 0.001187825 0.000550396 0.000550396 Case 8 0.000580341 0.001336675 0.000580341 0.000580341 Case 9 0.059404645 0.121309285 0.059404645 0.059404645 Case 10 0.0624699 0.134818114 0.0624699 0.0624699 Case 11 0.059404645 0.121309285 0.059404645 0.059404645 Case 12 0.0624699 0.134818114 0.0624699 0.0624699 Case 13 0.000580341 0.001336675 0.000580341 0.000580341 Case 14 0.000613732 0.001528176 0.000613732 0.000613732 Case 15 0.000580341 0.001336675 0.000580341 0.000580341 Case 16 0.000613732 0.001528176 0.000613732 0.000613732 Case 17 0.065868698 0.151712615 0.065868698 0.065868698 Case 18 0.000651201 0.001783724 0.000651201 0.000651201


Post-flight Absolute Value 0.101057318 0.101057318 0.110261081 0.110261081 0.110261081 0.110261081 0.121309285 0.121309285 0.001068804 0.001068804 0.001187825 0.001187825 0.001187825 0.001187825 0.001336675 0.001336675 0.121309285 0.121309285 0.134818114 0.134818114 0.121309285 0.121309285 0.134818114 0.134818114 0.001336675 0.001336675 0.001528176 0.001528176 0.001336675 0.001336675 0.001528176 0.001528176 0.151712615 0.151712615 0.001783724 0.001783724

184


Fuel Cells 32.5 0.603 0 0.25 Dish 9 -0.68 0 2.5 Dish 5 1.26 0 2.5

Load/Unload Device 0 0 0 0 Crew 1 --- 0.455 0.9 1.05 Crew 2 --- 0.455 -0.9 1.05




Empty Pressure Tank 48 -1.33 0.55 1.45 Empty Pressure Tank 48 -1.33 -0.55 1.45 Empty Pressure Tank 48 1.69 0 1.08 Empty Pressure Tank 0 0 0 0

Empty LOX Tank 12 1.3 0.55 0.54 Empty LOX Tank 12 1.3 -0.55 0.54 Empty LH2 Tank 17 -1.3 0.55 0.54 Empty LH2 Tank 17 -1.3 -0.55 0.54 Compressed Gas 34 - - -

LOX Fuel 617 - - - LH2 Fuel 103 - - -

Full Pressure Tank 56 -1.33 0.55 1.45 Full Pressure Tank 56 -1.33 -0.55 1.45 Full Pressure Tank 56 1.69 0 1.08 Full Pressure Tank 0 0 0 0

Full LOX Tank 351.4285714 1.3 0.55 0.54 Full LOX Tank 351.4285714 -1.3 -0.55 0.54 Full LH2 Tank 58.57142857 -1.3 0.55 0.54 Full LH2 Tank 58.57142857 1.3 -0.55 0.54

185


Value Case 1 0.023720689 0.038639399 0.023720689 0.023720689 Case 2 0.003623796 0.00071949 0.003623796 0.003623796 Case 3 0.003623796 0.00071949 0.003623796 0.003623796 Case 4 -0.018437602 -0.044799599 -0.018437602 0.018437602 Case 5 0.026039427 0.046382766 0.026039427 0.026039427 Case 6 0.003996232 0.000879733 0.003996232 0.003996232 Case 7 0.003996232 0.000879733 0.003996232 0.003996232 Case 8 -0.020435175 -0.056027569 -0.020435175 0.020435175 Case 9 -0.018437602 -0.044799599 -0.018437602 0.018437602 Case 10 -0.042766358 -0.10045657 -0.042766358 0.042766358 Case 11 -0.018437602 -0.044799599 -0.018437602 0.018437602 Case 12 -0.042766358 -0.10045657 -0.042766358 0.042766358 Case 13 -0.020435175 -0.056027569 -0.020435175 0.020435175 Case 14 -0.047665139 -0.129240688 -0.047665139 0.047665139 Case 15 -0.020435175 -0.056027569 -0.020435175 0.020435175 Case 16 -0.047665139 -0.129240688 -0.047665139 0.047665139 Case 17 -0.06973095 -0.170062657 -0.06973095 0.06973095 Case 18 -0.078203726 -0.226939799 -0.078203726 0.078203726

Post-flight Absolute Value 0.038639399 0.038639399 0.023721 0.038639

0.00071949 0.00071949 0.042095 0.08197 0.00071949 0.00071949 0.042095 0.08197

-0.044799599 0.044799599 0.018438 0.0448 0.046382766 0.046382766 0.026039 0.046383 0.000879733 0.000879733 0.046421 0.100227 0.000879733 0.000879733 0.046421 0.100227

-0.056027569 0.056027569 0.020435 0.056028 -0.044799599 0.044799599 0.089888 0.185841 -0.10045657 0.10045657 0.062991 0.141902

-0.044799599 0.044799599 0.089888 0.185841 -0.10045657 0.10045657 0.062991 0.141902

-0.056027569 0.056027569 0.099626 0.232418 -0.129240688 0.129240688 0.070207 0.182561 -0.056027569 0.056027569 0.099626 0.232418 -0.129240688 0.129240688 0.070207 0.182561 -0.170062657 0.170062657 0.069731 0.170063 -0.226939799 0.226939799 0.078204 0.22694

186




Value Case 1 0 0 0 0 Case 2 0.04193849 0.081967213 0.04193849 0.04193849 Case 3 -0.04193849 -0.081967213 -0.04193849 0.04193849 Case 4 0 0 0 0 Case 5 0 0 0 0 Case 6 0.046248715 0.100222717 0.046248715 0.046248715 Case 7 -0.046248715 -0.100222717 -0.046248715 0.046248715 Case 8 0 0 0 0 Case 9 0.08797654 0.180360721 0.08797654 0.08797654 Case 10 0.046248715 0.100222717 0.046248715 0.046248715 Case 11 -0.08797654 -0.180360721 -0.08797654 0.08797654 Case 12 -0.046248715 -0.100222717 -0.046248715 0.046248715 Case 13 0.097508126 0.22556391 0.097508126 0.097508126 Case 14 0.051546392 0.128939828 0.051546392 0.051546392 Case 15 -0.097508126 -0.22556391 -0.097508126 0.097508126 Case 16 -0.051546392 -0.128939828 -0.051546392 0.051546392 Case 17 0 0 0 0 Case 18 0 0 0 0


Post-flight Absolute Value 0 0

0.081967213 0.081967213 -0.081967213 0.081967213

0 0 0 0

0.100222717 0.100222717 -0.100222717 0.100222717

0 0 0.180360721 0.180360721 0.100222717 0.100222717

-0.180360721 0.180360721 -0.100222717 0.100222717

0.22556391 0.22556391 0.128939828 0.128939828 -0.22556391 0.22556391

-0.128939828 0.128939828 0 0 0 0

187

A.1.4 Materials Study (Jarred Young)

In order to design a mass-e�cient vehicle, the right material had to be chosen. There werethree overall factors that would play into the selection of our overall material: overall strength,density, and ease of machinability. To that e�ect, four materials were then selected to choosefrom. Their identities and properties are below:

Among these four metals, Aluminum alloy was chosen based on its relative strength to itsoverall density, making our craft more mass e�ective as a whole. Aluminum is also easier tomachine in most cases the other metals listed above. However, the choice of Aluminum Alloywould come to the properties of each alloy.

Out of these choices of aluminum alloys, Aluminum 6061-T6 was chosen, as it has the bestoverall characteristics for our vehicle. The alloy has a density of 2700 kg/m3, with a relativelyhigh yield strength of 310 MPa, and an Elastic Modulus of 68.9 GPa. Combined with itsresistance to fatigue, this material was chosen to be our overall material for the vehicle.

188

A.1.5 Moment of Inertia Analysis (Jarred Young)

189

190

191

192

A.1.6 Delta V Requirements for Multi-Crater Missions (Adam Kirk)

One study was done to look at more complex missions then just simply heading to one siteand then back to base. The purpose of the study was to determine the feasibility of going totwo di�erent sites in a single mission from a delta V standpoint.

To do this, the positions of �ve di�erent, important sites relative to each other were obtainedand distances between each of the �ve sites were calculated. The sites for the analysis werethe base at the rim of Shackleton Crater, Shackleton, Shoemaker, de Gerlache, and Malapert.After calculating the distance between each of these sites, the amount of delta V required foreach distance was calculated based on the ballistic hop equations found in Appendix 2.

The results of the analysis are:

Based on this analysis, it was found that the delta V requirement for even the shortest multi-crater mission (Base -> Shackleton -> de Gerlache -> Base) adds up to a ΔV requirement of1930 m/sec. This is greater than the longest mission the vehicle was designed for, which is toShoemaker and back to the base. For this reason, it was found that the vehicle would likelyonly be useful for missions consisting of 2 hops.

193

A.1.7 Glide vs Hop Trade Study (Adam Kirk)

Early in the design process, a quick study was done to compare the amount of ΔV requiredfor glide versus hopping to possible nearby mission sites. The four mission sites of Shackle-ton, Shoemaker, de Gerlache, and Malapert were used as examples. ΔV requirements werecalculated for each of these sites using both hopping and gliding trajectories to compare theamount of ΔV saved by using hops. For glides, it was assumed that the vehicle would glideto the edge of the crater, but then would have to hop down into the center for the elevationchange. The results of the study are shown below:

From this study, it was determined that the cost of gliding was hundreds of m/sec over ballistictrajectories for any of the sites nearby Shackleton Crater. It was felt that the large amountof extra fuel that would be necessary for this type of trajectory eliminated it as a possibilitysince minimizing weight was a priority.

194

A.2 MATLAB Scripts

A.2.1 Delta V Analysis (Adam Kirk)

195

196

A.2.2 Multi-Crater Analysis (Adam Kirk)

197

A.2.3 Range vs. Vehicle Mass Trade study (Adam Kirk, Neal Vasilak)

198

199

A.2.4 Tank Sizing Scripts (Nitin Sydney)

200

201

202

A.2.5 Thermal Scripts (Amirhadi Ekrami)

A.2.6 Moment of Inertia Scripts (Jarred Young)

203

204

A.2.7 Payload Bay Sizing (Fazle Siddique)

205

A.3 Payload Database (Fazle Siddique)

206

207

208

A.4 Ballistic Hop and Glide Equation Derivations (Adam Kirk)

The following equations for ballistic hops and glides assume that a vehicle is located on an �atbody without an atmosphere. Thus, drag e�ects and the curvature of the body are not takeninto account.

A.4.1 Ballistic Hop on Flat Airless Body

Following basic kinematics:

gVv −=& 2

2

1gttVh v −= gVt vflt 2=

where Vv is the vertical component of the velocity, h is the vertical height of the vehicle, t istime, and g is the gravitational acceleration induced by the body. Because there is no drag,the horizontal component of the velocity, Vh, will remain constant. Thus, it follows that:

g

VVtVd vh

flth 2== γcosVVh = γsinVVv =

( )γ2sin2

g

Vd =

209

where d is the horizontal distance and γ is the launch angle of the vehicle. Figure ?? visuallyshows the relationship between these variables.

For a ballistic trajectory, the maximum horizontal distance is achieved with a launch angle of45 degrees. Thus, the change in velocity for an optimum ballistic hop can be found with thefollowing equations:

°= 45optγ gdVg

Vd =⇒=

2

max gdVV 22 ==∆

A.4.2 Ballistic Hop on Flat Airless Body with Elevation Change

This case is the same as done in the previous section, except now it is assumed that the vehicleis landing at some elevation h2 relative to the starting point. This means that the optimumlaunch angle is no longer 45 degrees. It is then also true that the vertical components of thevelocity change will not be the same at the starting and ending points. If vv1 is the initialvertical velocity component, and vv2 is the landing vertical velocity component, then:

21 2

1gttvh v −=

g

vt v

peak1= peakv

vpeak ghv

g

vh 2

2

11

21 =⇒=

Then, to get vv2 in terms of hpeak and h2:

fallv gtv −= and 2

2

1fallpeak gthh −= so

g

vhh v

peak

22

2 2

1−=

( )22hh

gt peakfall −= ( )22 2 hhgv peakv −=

210

Then, to �nd the horizontal component of velocity:

( )

−+=+= )(2

21 hh

gg

vvttvd peak

vhfallpeakh

)(22 2hhh

gdv

peakpeak

h −+=

With each of the velocity equations, one can then calculate the totalΔV required for a selectedhpeak through the equation:

( )22

221

2vhvhtot vvvvV +++=∆

In order to �nd the optimum trajectory, the peak height must be iterated until a minimumΔV is found.

A.4.3 Propulsive Glide on Flat Airless Body

First, assume that the vehicle glides at some horizontal velocity V, so that:

VVh 2=∆

Then, using the time of �ight, t�t, it follows:

Vdt flt =

V

gdgtV fltv ==∆

V

gdVVVV hvtot +=∆+∆=∆ 2

The minimum V can be found by taking the derivative with respect to V and �nding a thepoint of zero slope. Doing this, the optimum V is found to be:

2

gdVopt =∆ so gdVtot 22=∆

211

A.5 Dummy Data

A.6 Outreach

A.6.1 Hours List

212

Figure 121: Outreach Hours List

A.7 Additional References

Hap Ehlers. �Shuttle Orbiter / Cargo Standard Interfaces ICD 2-19001 Revision-L.�

http://www.unitedspacealliance.com/icd/.

Markley, F. Landis. "Fast Quaternion Attitude Estimation from Two Vector Measurements."October 2001. NASA Technical Reports Server. 5 May 2009 <http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20010068636_2001110260.pdf>.

NASA. "NASA's Lunar Communications & Network Architecture." Technology ExchangeConference. 2007.

Amy Ross, �Lunar Rover Vehicle Mockup Advanced Space Suit Ingress Egress Test,� 2002.

�Polymer Data Files.� Tangram Technology Ltd. 25 Nov 2005. Tangram Co. UK. 08 May2009. <http://www. tangram.co.uk/TI-Polymer-Introduction.html>.

�TexLoc Refractive Index of Polymers.� The TexLoc Closet. 09 Oct 2008. Parker-TexLoc. 08May 2009. <http:// www.texloc.com/closet/cl_refractiveindex.html>.

�Material Property Data: Thermoplastic.� MatWeb. Automation Creations, Inc. 08 May2009. <http://www. matweb.com/Search/MaterialGroupSearch.aspx?GroupID=12>.

Ware, J., et al. �Design and Testing of Improved Spacesuit Shielding Components.� Society ofAutomotive Engineers, Inc. 2002-01-330. 2002. pp 6. Available online: <http://www.osti.gov/bridge/servlets/purl/825125-y1Hh33/native/825125.pdf>.

213

Rais-Rohani, M. �On Structural Design of a Mobile Lunar Habitat With Multi-Layered En-vironmental Shielding.� NASA. NASA-CR-2005-213845. April 2005. pp 9. Available online:<http://marsjournal.org/contents/ 2006/0004/�les/Rais2005.pdf>.

http://space�ightsystems.grc.nasa.gov/Advanced/Capabilities/Energy/

214