Lunar Exploration Transportation System (LETS)

MAE 491 / 4922008 IPT Design Competition

Instructors: Dr. P.J. Benfield and Dr. Matt Turner

Team Frankenstein

Final Review Presentation4/29/08

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

Overview• Mission Statement• The Need• The Solution• Performance• Schedule• Operations• Structures• GN&C

• Communications• Payload• Power• Thermal• Risk Management• Conclusions• Questions

Mission Statement

• 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

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

The Solution

• Lander/Rover

• Penetrators

• RTG

Cyclops

PerformanceCDD Requirement Requirement Assessment Remark

Landed Mass 932.8 kg Exceeds Actually 810 kg

Survive Lunar Cruise 28 days Meets Capable of surviving lunar cruise exceeding 28 days

Operational Period 1 year Meets TRL 9 materials will remain functional beyond 1 year

Sample Lunar Surface 15 dark Exceeds Mobility allows roving to as many sites as is needed

CommunicationSend and Receive

(real time) Exceeds Capable of sending data at 150 Mbps

Landing Parameters 12º slope within 100 m ExceedsSix wheel rocker bogey system allows landing on slopes

greater than 12 degrees

Survive Launch of 6 G's 6 G's Exceeds Cyclops structure will handle g-loads exceeding 6 g’s

TechnologyRequirements TRL9 Meets Materials used are TRL 9

Power Requirements Store Power in Dark

Conditions Exceeds RTG can provide the power needed during dark conditions

Thermal ConditionsSurvive Temperature

Changes ExceedsMaterials used will withstand temperatures exceeding the

50K to 380K range

Sample Return Vehicle Sample Return (Goal) Meets Exceeds the sample return expectations

MobileRoving/Real-Time

Mobility Exceeds 6 wheel rocker bogey allows roving in real-time

Schedule

Operations

Cyclops

Penetrators 2.5km

1.6km

1. 5km

2. Deploy Penetrators

3-4. Decent

5. Land

6. Release Propulsion System

7. Rove To Edge of Crater

Operations• Launch - September 30, 2012• Arrive at moon - October 6, 2012• Operations start 5km from lunar surface• October 8, 2012

– Decent • Shoot 15 penetrators into Shackleton Crater for dark region sampling

– Landing• Drop off “single site box” to accomplish single site goals

• October 9, 2012– Rove to rim of Shackleton Crater

• October 11 - 18, 2012– Receive all data from penetrators

• October 19, 2012– Relay all data from penetrators to LRO for transmission to Mission

Control• 5 orbits needed

Operations• October 22, 2012 – March 4, 2013

– Rove to, collect and relay data from 29 lighted sites• March 5 – March 7, 2013

– Rove to, collect sample, and launch SRV• March 8 – July 22

– Rove to, collect and relay data from Lighted sites 30 - 59• July 23 – 25

– Rove to rim of Shackleton crater• July 26 – September 27

– Rove to, collect and relay data from Dark Sites (if penetrators fail)

• September 30, 2013– System Shut Down

Structures• 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

Penetrator Ring Platform

Attitude Control System

Cyclops

Structures• 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 90 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.

Before Deployment

After Deployment

Structures

• Maxon 60mm EC 60 x 6

• Nominal torque 830 mNm

• Maxon 45mm EC 45 x 8

• Nominal torque 310 mNm

Wheel Motors Steering Motors

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

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

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

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: 10 Kg– Dimensions: 155x68.5x35.5 cm– Power: Solar Panel

Payload• Gas Chromatograph Mass Spectrometer (GCMS)

– Performs atmospheric and organic analysis of the lunar surface – Mass: 19 Kg– Dimensions: 10x10x8 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: 110X10X10 cm– Power: Rover

Payload• Penetrators (Deep Space 2 )

– Mission’s main source of data acquisition in the permanent dark regions

– Mass (15 Penetrators): 53.58 Kg – Dimensions: 13.6Dx10L 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: 23.5x16.3x15.5 cm– Power: Rover

Power• RTG

– TRL9– Constant power supply– Thermal output can be utilized for

thermal systems• Lithium-Ion Batteries

– Commercially available– Easily customizable– Rechargeable

• Solar – Used for Single Site Box– Conventional– Increasingly efficient in well light

areas

POWER SUBSYSTEM

Type (solar, battery, RTG) Solar, Lithium-ion, RTG

Total mass 47.63 kg

Total power required 643.525 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 1

RTG Mass/RTG 40 kg

PowerPower Analysis

Component Subcomponents Consumption (W)

Mobility   342.625

SRV   25

GN&C   115.5

Payload   34.6

Communications   70.8

Thermal   55

Operations   0

Power Supply   865

RTG 400

  Li-ion Battery 455

Solar Cell 10 (Not in Total)

Minimum Totals   643.525

Contingency Supply 33% 212.36325

Total   855

• Total Power Required– 643.525 W

• Peak Power– Mobility

• 342.625 W– Data Collection/Transfer

• 276.2 W– Single Sight Box

• 7.8 W• RTG

– 400 W • Lithium-Ion Batteries

– 455 W for both• Solar Cells

– 10 W (SSB - Not included in Total)• 33% Contingency Power• Total Power Supplied to Lander

– 855 W

Thermal

• Cyclops uses three types– Heat transfer pipes– Paraffin heat switches

• Radiator heat switches• Diaphragm heat

switches

– Multi-Layer Insulation

Thermal• Two standard types of switches are used as a

redundant check to prevent over heating

Thermal

• Heat is well controlled– MLI has low heat

absorbance– Heat switches allow close

tolerance controlMass (kg)

Heat pipes 4.62

Heat Switches 3

MLI 3.78

Total 11.4

Risk Management5 1

4 9 2 3

3 8 5 4

2 6,7 10

1

1 2 3 4 5

LIKELIHOOD

CONSEQUENCES

# Mission Criticality Level Risk1 high Penatrator Failure - Penatrators fail upon impact or do not launch at all2 medium RTG Ovherheat - RTG puts out too much heat for the system to handle3 high Thermal Shutdown - the system gets overheats and shuts down4 medium Mobility - the drive system of the lander fails5 medium Camera Failure - the camera breaks and does not transmit images6 medium Structure Failure - the structure collapes7 medium Navigation - the navigation system fails8 low SRV Failure - the SRV fails to launch9 low Single Site Box - the box fails during the year

10 medium Communications - loss of all communications with the rover

Likelihood1 improbable23 probable45 definite

Consequences1 the mission can still be completed23 mission operates at limited capacity45 total mission failure

Likelihood1 improbable23 probable45 definite

Conclusions

• “There’s no place this thing can’t go”

• If penetrators fail, remaining mission will not be compromised

• Reliable multi-faceted design

Questions