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

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 - CDD Requirement Requirement Assessment
Remark
Landed Mass
932.8 kg
Actually 785 kg (16% mass margin)
Survive Lunar Cruise
28 days
Capable of surviving lunar cruise exceeding 28 days
Operational Period
1 year
TRL 9 materials will remain functional at least 1 year
Sample Lunar Surface
15 dark
5 lighted
Mobility allows roving to as many sites as is needed
Communication
Send and Receive
(real time)
Capable of sending data at 150 Mbps
Landing Parameters
12º slope within 100 m
Six wheel rocker bogey system allows landing on slopes greater than 12 degrees
Survive Launch of 6 G's
6 G's
Cyclops structure will handle g-loads exceeding 6 g’s
Technology
Requirements
TRL9
Materials used are TRL 9
Power Requirements
Store Power in Dark
Conditions
RTG can provide the power needed during dark conditions
Thermal Conditions
Survive Temperature
Changes
Materials used will withstand temperatures exceeding the 50K to 380K range
Sample Return Vehicle
Sample Return (Goal)
Exceeds the sample return expectations
Mobile
Roving/Real-Time
Mobility
6 wheel rocker bogie allows roving in real-time

8. Schedule 8

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 9

10. Structures System Specifications (Auxiliary Systems)
Penetrator Ring Platform
Outer Diameter 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 m Outer Diameter
Aluminum Frame (6061 T6)
Single Site Box
Max Box Dimensions x x m
Integrated Sample Return Vehicle
Attitude Control System
Cyclops
Penetrator Ring Platform

11. Structures System Specifications (Main) Main Chassis
Before Deployment
System Specifications (Main)
Main Chassis
Dimensions – 1.54 x 1.54 x 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 14
Camera
(SSI) Dimensions x x m
Scoop Arm
Max Reach m
After Deployment

12. Structures Wheel Motors Steering Motors Maxon 60mm EC 60 x 6
Nominal torque 830 mNm
Maxon 45mm EC 45 x 14
Nominal torque 310 mNm

13. GN&C Decent/Landing Post 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 Processor
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 Penetrators 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

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
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

18. Payload Penetrators (Deep Space 2 )
Mission’s main source of data acquisition in the permanent dark regions
Mass (15 Penetrators): kg
Dimensions: 13.6D x 10L cm
Power: 2 Lithium Ion Batteries Each
Miniature Thermal Emission Spectrometer (Mini-TES)
Objective 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

19. Power RTG Lithium-Ion Batteries Solar Cell TRL9 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
POWER SUBSYSTEM
Type (solar, battery, RTG)
Solar, Lithium-ion, RTG
Total mass
63.62 kg
Total power required W
Number of solar arrays 1
Solar array mass/solar array
1.13 kg
Solar array area/solar array
0.12 square meter
Number of batteries 2
Battery mass/battery
3.25 kg
Number of RTGs
RTG Mass/RTG
56 kg 19

20. Power Power Analysis Total Power Required 431.35 W Peak Power Roving
Mobility, GN&C and Thermal 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
Power Analysis
Component
Subcomponents
Consumption (W)
Mobility
112.15
SRV 1
GN&C
115.5
Payload
76.9
Communications
70.8
Thermal 55
Operations
Totals
431.35
Power Supply
RTG
300
Li-ion Battery
975W-h
Solar Cell
10W 20

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 of Rover / Mass of Total System (%) = 65%
Mass Allocations
Mass of Total System
Mass of Rover
System
Mass (kg)
Percent Mass
GN&C
23.5 3%
Payload 90
11.5%
Communications 20
2.5%
Thermal
11.4
1.5%
Power Supply 64
8.2%
Structures
173.1
22.1%
SRV
153
19.5%
Mobility
215
27.4%
SSB 35
4.5%
Total Mass
785 kg
100%
System
Mass (kg)
Percent Mass
GN&C
23.5
4.6%
Payload
36.4
7.1%
Communications 20
3.9%
Thermal
11.4
2.9%
Power Supply 63
12.3%
Structures
138.1
27%
Mobility
215
42.1%
Total Mass
511 kg
100%
Mass of Rover / Mass of Total System (%) = 65% 25

26. Figures of Merit Figure of Merit Goal Design
Number of surface objectives accomplished
15 samples in permanent dark
5 samples in lighted terrain
15 penetrators taking samples in permanent dark and 5 lighted samples taken by Cyclops
Percentage of mass allocated to payload
Higher is better
11.5%
Ratio of objectives (SMD to ESMD) validation
2 to 1
Efficiency of getting data in stakeholders hands vs. capability of mission
95%
Percentage of mass allocated to power system
Lower is better
8.2%
Ratio of off-the-shelf hardware to new development hardware
85% 26

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