Final Report on LOW Design

Maximizing ScienceWhile

Minimizing Single Point Failure

Project ManagementTeam Eclipse

Project Office Eddie Kiessling

Systems Engineer Jay Gala

Structures Nathan Coffee

GN & C Brandon York

GN & C Joseph Sandlin

Operations Brett Guin

Thermal Kathryn Kirsh

Payload Operations Brent Newson

Power Christopher Goes

Sample Return Vehicle Julien Gobeaut

Sample Return Vehicle Ghislain Pelieu

Technical Editor Michael Bryan

Team Objectives

• Develop Teamwork– Work as a team in an engineering

environment

• Systems Integration

• Time management

• Develop communication skills

Final Concept Drawing

Video

Figures of Merit Table

Figure of Merit Goal Design

Number of surface objectives accomplished

15 samples in permanent dark and 5 samples in lighted

terrain

15 samples in permanent dark and 5 samples in lighted

terrain

Percentage of mass allocated to payload

Higher is better 33%

Ratio of objectives (SMD to ESMD) validation

2 to 1 4:1

Efficiency of getting data in stakeholders hands vs. capability of mission

Higher is better 58%

Percentage of mass allocated to power system

Lower is better 26%

Ratio of off-the-shelf hardware to new

development hardwareHigher is better

All subsystems have TRL 9 equipment

Critical Parameters of LOWParameter Units Notes

Overall Vehicle

Mission Duration Days 331

Total Mass kg 997.4

Number of Sites Visited Sites 20 Total (5 Light / 15 Dark)

Single Site Goal Mass kg 23

Payload Subsystem

Total Mass kg 324

SMD Mass kg 260

ESMD Mass kg 64

Payload Percentage of Total Mass % 32.48

Power Subsystem

Type N/A Solar Cells & Lithium Ion Batteries

Total Power Mass kg 263.698

Total Power Required We 200 We

Number of Solar Arrays N/A 1

Solar Array Mass / Solar Array kg 18.203

Solar Array Area / Solar Array m2 2.217

Number of Batteries N/A 24

Batter Mass / Battery kg 9.824

Critical Parameters of LOWParameter Units Notes

Structure Subsystem

Total Mass kg 140

Maximum G Load G 5

Thermal Subsystem

Temperature of Rover (Cold Case) K 160 – 250

Temperature of Rover (Hot Case) K 226 – 277

Total Mass Kg 40

Passive / Active System N/A Active

GN&C Subsystem

Total Mass kg 42.164

Accuracy % 95

Power Required W

Communication Subsystem

Total Mass kg 5

Type N/A S – Band Transmitter

Bandwith MHz 1700 – 2300

Power Required W 30

Data Rate Bits Per Minute (bpm) 4320000

Critical Parameters of LOW

Parameter Units Notes

Mobility System

Range of Velocity m / hr 5 – 6

Total Mass kg 100

Power Allocation Table

Average Load Landing Initializing Driving Science Communications Sleep EOM & SRV

Thermal 40 40 40 40 40 10 2

GN&C 50 2 23 15 10 1 10

Power 10 10 10 10 10 1 10

C&DH 2 20 5 20 20 1 40

Communications 2 30 2 2 30 1 2

Propulsion 5 0 0 0 0 0 0

Mobility 4 10 100 10 10 0 10

Mechanisms 4 2 2 2 2 1 2

Payload 2 2 2 75 2 1 10

SRV 1 1 1 1 1 1 10

Total 120 117 185 175 125 17 96

CDD Requirements Atlas V-401 EPF with landed mass of 997.4 kg Propulsion system dry mass is 64.6 kg First mission at polar location Capability to land at other lunar locations Minimize Cost Launch Date is Sept. 30, 2012 Capability to move on lunar surface Survive for 1 year on the lunar surface Survive the proposed concept of operations Must meet SMD and ESMD objectives Land within a precision of ± 100m 3σ Provide guidance, navigation, and control beginning at 5 km above lunar

surface Capable of landing at a slope of 12 degrees Designed for G-loads during lunar landing Design to withstand g-loads with respect to stiffness only

Payload Subsystem

• Stereo Imaging• Belly Cam• GN&C

Isometric View of LOW

Payload Subsystem

• Upon landing, the LOW prepares for single-site goals and multi-site goals.

• A drop-box is prepared for single-site goals utilizing various instruments measuring the following parameters:– Lighting conditions– Micrometeorite flux– Electrostatic dust levitation conditions

• Meets CDD Single Site Goals• The LOW’s stereo imaging systems’ mast will

permanently raise itself to its functional position.

Payload Subsystem• At each site starting at site one, scientific equipment will

perform multi-site goals utilizing various instruments performing the following operations:– Regolith sample collection– Geotechnical properties– Regolith composition– Geological characterization– Magnetic susceptibility– Surface temperatures– Alpha particle and gamma ray emissions– Image acquisition

• Meets and exceeds CDD Multi-site instrument package goals• Maximized scientific equipment by allotting 325 kg to Payload• 32 Individual pieces of equipment

Payload SubsystemTop View of Payload Layout

Payload Subsystem

• Sample Return Vehicle– Utilizes liquid propellant

• NTO / MMH

– Separate GN&C system• 2 Ring Laser Gyros• Pressure Sensors• Air Speed Sensors• Angle of Attack Sensors

– Launches directly from LOW• Site of Launch: Shackleton Crater

Payload Subsystem

• At the final site (Shackleton Crater), the SRV will be ready to launch.

• An arm will load the SRV with regolith samples collected during the course of the mission.

Structures Subsystem

• 5 G’s Maximum• Semi – Hard Landing

Utilizing Crush Pads– Adapted from Mars Viking

Lander

• Experiences impact velocity of 8 m / s

• Minimum Factor of Safety of 1.5– Verified by Von Mises

Stress Criterion

• Four Leg Configuration

Structures Subsystem

• Landing Phase– Four Unloading

Ramps• Primary System

– Explosive Bolts

• Secondary System– Robotic Arm Assist

Guidance, Navigation, & Control Subsystem

• Descent (Terminal Descent Phase – TDP)– Attitude Controllers (ACS)

• IMU (Northrop Grumman LN-200 Fiber Optic)• Star Tracker (Goodrich HD-1003)• Sun Sensors (Optical Energy Technology Model 0.5)• Radar Altimeter (Honeywell HG8500)• Actuators

– Thrusters/Main Engine (AeroJet MR-106 and MR-80B)– Reaction Wheels (Ithaco Type-A)

– Optical (OSS)• DSMAC (Digital Scene Matching Area Correlator)• LiDAR (Light Detection and Ranging) (MDA-Optech)

– Guidance Computer (BAE Systems RAD750)

Guidance, Navigation, & Control Subsystem

• Experimentation and Traveling (Lunar Excursion Phase – LEP)– 2 Panoramic Cameras (PanCams)– 4 Hazard Cameras (HazCams)– 1 Navigational Camera (NavCam)– 1 Belly Camera (BellyCam)– Also uses the IMU, Tilt Sensors, and the Guidance

Computer

Thermal Subsystem

• To account for the diverse temperature range (50K – 400K) of the moon, the LOW utilizes a active thermal system with the following components:– Radiators– Heaters– Heat Pipes– Multi – Layer Insulation (MLI)

Power Subsystem

• Use of Nuclear Power was avoided– Nuclear power conflicts with the 2012 launch

• Requires a 5 year lead time– Pu-238 is scarce

Power Subsystem• Power Source

– Solar Arrays• Area of 2.217m2

• Silicon• Mass of 18.203 kg• Solar Power Required 455 We

• Power Storage– Lithium Ion Batteries

• 24 Batteries • Mass of 9.824 kg per battery• 45 kWe-hr Capacity

• During the Dark– 8 days of power at 200 We average– 6 days of sleep phase at 20 We

Communications Subsystem

• LOW uses S-band transmitters for communication– Relays data through LRO to Earth at DLS– Relays data to Earth at DLS

• Data transfer occurs after each site– Requires approximately 5.5 days to transfer 300 MB

of experimental data for each site• DLS of LRO for approximately 8.5 minutes per orbit• Orbit Time of 113 minutes• Data Rate of 72000 bits per second (bps)

Communications Subsystem

• Communication and Data– Gives and receives telemetry data when

in LOS to LRO and Earth

– Stores all data on 2 different 32 GB non-volatile memory

• Has no moving parts• Redundancy

Sample Return Vehicle (SRV)

SRV Mass Balance

SRV GN&C

• Control of the rocket– Control in pitch & yaw– 2 actuators– Sensor: 1 IMU

• Control of the re-entry capsule– Control on 3 axis– 6 thrusters– Sensor:

1 IMU

3 pressure sensors on the heat shield

Actuator

Nozzle

SRV Trajectory

Conclusion

• Maximize Scientific Payload– LOW has the potential to visit extra sites– LOW can perform more in-depth experiments

• Minimize Single Point Failure– Comprised mostly of TRL 9 technology– Conservative ConOps Schedule– No lag time in data communication– Complete Redundancy in GN&C

Questions?

Landing Site

Landing Landing PhaseBattery Power

GN&CLanding Legs

Initializing Initializing PhaseCom. Sys. “OK”

Solar PowerScience BoxCameras on

1st Science Site Science PhaseExperimentsCom. Phase

Data Transfer1 day

1st day

1st day

Light Sites

Travel to Site Traveling PhaseMobility

Solar PowerCharge Batteries

Com. PhaseData Transfer Repeat

forLight

Science Site Science PhaseExperimentsCom. Phase

Data Transfer

4 days

1 day

Dark Sites

Travel to Dark Traveling PhaseMobility

Battery Power

Science Site Science PhaseExperimentsCom. Phase

Data Transfer

Travel to Light Traveling PhaseMobility

Battery PowerCom. Phase

Data Transfer

Await Sunlight Sleep PhaseMinimal Thermal

And power

Repeatfor

Dark

4 days

1 day

3 days

3 days

Shackleton Crater

Travel to EoMTraveling Phase

MobilityBattery PowerSolar Power

Shackleton Science PhaseExperimentsCom. Phase

Data TransferEoM & SRV Phase

Launch SRV

3 Weeks

Radiator Temperature

255

275

295

315

335

355

375

395

415

0.5 0.6 0.7 0.8 0.9 1

Emissivity

Tem

p o

f R

adia

tor

(K)

0.5 kW

1.0 kW

1.5 kW

2.0 kW

Radiator Temperature vs. Area

150

200

250

300

350

400

450

0 5 10 15 20 25

Area (sq. meters)

Ra

dia

tor

Te

mp

(K

)

0.5 kW

1.0 kW

1.5 kW

2.0 kW

Cold Case

Cold Case

30

80

130

180

0.04 0.09 0.14 0.19 0.24 0.29 0.34

Emissivity

T o

f R

ove

r (K

) 50 W

100 W

150 W

200 W

0 W

Hot CaseHot Case

208

218

228

238

248

0.04 0.09 0.14 0.19 0.24 0.29 0.34

Emissivity

T of

Rov

er (K

)

50 W

100 W

150 W

200 W

0 W

Expected Temperature Ranges

Equipment Temperature Range (C) Batteries -10 to 20 Electronics 0 to 40 Structures -46 to 65 Cameras -55 to 295 Solar Arrays -100 to 100 Power Box Baseplates -10 to 50 C& DH Box Baseplates -20 to 60 Antennas -100 to 100 Mobility -80 to 200