Final Report on LOW Design Maximizing Science While Minimizing Single Point Failure.
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Transcript of Final Report on LOW Design Maximizing Science While Minimizing Single Point Failure.
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Final Report on LOW Design
Maximizing ScienceWhile
Minimizing Single Point Failure
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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
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Team Objectives
• Develop Teamwork– Work as a team in an engineering
environment
• Systems Integration
• Time management
• Develop communication skills
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Final Concept Drawing
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Video
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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
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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
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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
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Critical Parameters of LOW
Parameter Units Notes
Mobility System
Range of Velocity m / hr 5 – 6
Total Mass kg 100
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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
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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
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Payload Subsystem
• Stereo Imaging• Belly Cam• GN&C
Isometric View of LOW
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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.
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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
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Payload SubsystemTop View of Payload Layout
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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
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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.
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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
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Structures Subsystem
• Landing Phase– Four Unloading
Ramps• Primary System
– Explosive Bolts
• Secondary System– Robotic Arm Assist
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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)
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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
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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)
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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
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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
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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)
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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
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Sample Return Vehicle (SRV)
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SRV Mass Balance
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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
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SRV Trajectory
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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
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Questions?
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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
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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
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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
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Shackleton Crater
Travel to EoMTraveling Phase
MobilityBattery PowerSolar Power
Shackleton Science PhaseExperimentsCom. Phase
Data TransferEoM & SRV Phase
Launch SRV
3 Weeks
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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
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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
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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
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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
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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