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Lunar Exploration Transportation System (LETS)

Lunar Exploration Transportation System (LETS). MAE 491 / 492 2008 IPT Design Competition Instructors: Dr. P.J. Benfield and Dr. Matt Turner Team Frankenstein Final Review Presentation 4/29/08. Team Disciplines. The University of Alabama in Huntsville Team Leader: Matt Isbell

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Lunar Exploration Transportation System (LETS)

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  1. Lunar Exploration Transportation System (LETS) MAE 491 / 492 2008 IPT Design Competition Instructors: Dr. P.J. Benfield and Dr. Matt Turner Team Frankenstein Final Review Presentation 4/29/08

  2. Team Disciplines • The University of Alabama in Huntsville • Team Leader: Matt Isbell • Structures: Matthew Pinkston and Robert Baltz • Power: Tyler Smith • Systems Engineering: Kevin Dean • GN&C: Joseph Woodall • Thermal: Thomas Talty • Payload / Communications: Chris Brunton • Operations: Audra Ribordy • Southern University • Mobility: Chase Nelson and Eddie Miller • ESTACA • Sample Return: Kim Nguyen and Vincent Tolomio

  3. Overview • Payload • Power • Thermal • Risk Management • Mass Allocations • Figures of Merit • Conclusions • Questions • Mission Statement • The Need • The Solution • Performance • Schedule • Operations • Structures • GN&C • Communications

  4. Mission Statement Team Frankenstein’s mission is to provide NASA with a reliable and multi-faceted lander design that will provide the flexibility to conduct CDD requirements, scientific investigations, and technology validation tasks at different areas on the moon.

  5. The Need • Only 6% of lunar surface explored • Apollo missions • Only orbital visits since Apollo • Mobile lunar laboratory with return capabilities is vital to the exploration and understanding of the lunar surface • The lunar surface is an unexploited record of the history of the solar system • Sample polar sites and crater floors

  6. The Solution Cyclops • Lander/Rover • Penetrators • RTG

  7. Performance Exceeds - Meets -

  8. Schedule

  9. Operations 1. 5km 2. Deploy Penetrators 3-4. Decent 5. Land 6. Release Propulsion System 7. Rove To Edge of Crater Penetrators 2.5km Cyclops 1.6km

  10. Structures Attitude Control System • System Specifications (Auxiliary Systems) • Penetrator Ring Platform • Outer Diameter - 3.189 m • Aluminum Construction (6061 T6) • Mounted penetrators (spring released at a 4 degree dispersion angle) • Attitude Control • Main thrusters - MR 80B • Attitude Control Thrusters - MR 106 • Hydrazine Tank x 2 - 0.549 m Outer Diameter • Aluminum Frame (6061 T6) • Single Site Box • Max Box Dimensions - 1.54 x 0.688 x 0.356 m • Integrated Sample Return Vehicle Cyclops Penetrator Ring Platform

  11. Structures Before Deployment • System Specifications (Main) • Main Chassis • Dimensions – 1.54 x 1.54 x 0.356 m • Aluminum Frame (6061 T6) • Carbon composite exterior • MLI Insulation • 6 Wheel Passive Rocker Bogie Mobility System • Proven Transportation Platform (MER, Pathfinder) • 0.33 m Outer Diameter Wheels • Can navigate up to a 45 degree angle • Max speed of 75 m/hr • Aluminum construction (6061 T6) • Maxon EC 60 Brushless DC motor (60mm) x 6 • Maxon EC 45 Brushless DC motor (45mm) x 8 • Camera • (SSI) Dimensions - 0.305 x 0.203 x 0.152 m • Scoop Arm • Max Reach - 1.727 m After Deployment

  12. Structures Wheel Motors Steering Motors • Maxon 60mm EC 60 x 6 • Nominal torque 830 mNm • Maxon 45mm EC 45 x 8 • Nominal torque 310 mNm

  13. GN&C • Decent/Landing • A LIDAR system will be used to control, navigate, and stabilize while in descent • Post Landing • An operator at mission control will manually navigate lander/rover • A Surface Stereo Imager (SSI) periscopic, panoramic camera will be used to survey the lunar surface, provide range maps in support of sampling operations, and to make lunar dust cloud measurements

  14. GN&C • Descent Imaging • A Mars Descent Imager (MARDI) will be used to view both the penetrator dispersion and the landing/descent of the Cyclops • Processor • A BAE RAD750 will be used for all controls processing

  15. Communications • Rover • Parabolic Dish Reflector Antenna (PDRA) • T-712 Transmitter • Communication Bandwidth : X-band • Data Transmission Rate: 150 Mbps • Data Storage Capacity: 10 GB • Penetrators • Omnidirectional Antenna • Communication Bandwidth: S-band • Data Transmission Rate: 8 Kbps • Data Storage Capacity: 300 MB

  16. Communications/Payload • Single Site Box (SSB) • Determines lighting conditions every 2 hours for one year, micrometeorite flux, and assess electrostatic dust levitation • Omnidirectional Antenna • Communication Bandwidth: S-band • Data Transmission Rate: 8 Kbps • Data Storage Capacity: 1GB • Surface Stereo Imager (SSI) • Mass: 35 kg • Dimensions: 155 x 68.5 x 35.5 cm • Power: Solar Panel (7.8 W)

  17. Payload • Gas Chromatograph Mass Spectrometer (GCMS) • Performs atmospheric and organic analysis of the lunar surface • Mass: 19 kg • Dimensions: 10 x 10 x 8 cm • Power: Rover (60 W) • Surface Sampler Assembly (SSA) • Purpose is to acquire, process, and distribute samples from the moon’s surface to the GCMS • Mass: 15.5 kg • Dimensions: 110 x 10 x 10 cm • Power: Rover (3.5 W)

  18. Payload • Miniature Thermal Emission Spectrometer (Mini-TES) • Objective is to provide measurements of minerals and thermo physical properties on the moon • Mass: 2.4 kg • Dimensions: 23.5 x 16.3 x 15.5 cm • Power: Rover (5.6 W) • Penetrators (Deep Space 2 ) • Mission’s main source of data acquisition in the permanent dark regions • Mass (15 Penetrators): 53.58 kg • Dimensions: 13.6D x 10L cm • Power: 2 Lithium Ion Batteries Each (0.3 W)

  19. Power RTG TRL 9 Constant power supply Thermal output can be utilized for thermal systems Lithium-Ion Batteries Commercially available Easily customizable Rechargeable Solar Cell Used for Single Site Box Conventional Increasingly efficient in well light areas

  20. Power • Total Power Required • 431.35 W • Peak Power • Roving • Mobility, GN&C, and Thermal • 243.65 W • Data Collection/Transfer • All subsystems except mobility • 279.2 W • Single Sight Box/Sample Return Vehicle • 8.8 W • RTG (Cassini) • 300 W • Lithium-Ion Batteries • 975 W-h for both • Solar Cells • 10 W (SSB and SRV) • 7% RTG Contingency Power • Total Power Supplied to Lander • 300 W constant supply • 975 W-h for peak power outputs

  21. Thermal Cyclops uses three forms of heat control Heat transfer pipes Paraffin heat switches Radiator heat switches Diaphragm heat switches Multi-Layer Insulation

  22. Thermal • Two standard types of switches are used as a redundant check to prevent over heating

  23. Thermal • Heat is well controlled • MLI has low heat absorbance • Heat switches allow close tolerance control

  24. Risk Management LIKELIHOOD CONSEQUENCES

  25. Mass Allocations

  26. Figures of Merit

  27. Conclusions • “There’s no place this thing can’t go” • If penetrators fail, remaining mission will not be compromised • Reliable multi-faceted design

  28. Questions

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