Final Report on LOW Design Maximizing Science While Minimizing Single Point Failure

Project Management Team Eclipse Project OfficeEddie Kiessling Systems EngineerJay Gala StructuresNathan Coffee GN & CBrandon York GN & CJoseph Sandlin OperationsBrett Guin ThermalKathryn Kirsh Payload OperationsBrent Newson PowerChristopher Goes Sample Return VehicleJulien Gobeaut Sample Return VehicleGhislain Pelieu Technical EditorMichael 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 MeritGoalDesign Number of surface objectives accomplished 15 samples in permanent dark and 5 samples in lighted terrain Percentage of mass allocated to payload Higher is better33% Ratio of objectives (SMD to ESMD) validation 2 to 14:1 Efficiency of getting data in stakeholders hands vs. capability of mission Higher is better58% Percentage of mass allocated to power system Lower is better26% Ratio of off-the-shelf hardware to new development hardware Higher is better All subsystems have TRL 9 equipment

Critical Parameters of LOW ParameterUnitsNotes Overall Vehicle Mission DurationDays331 Total Masskg997.4 Number of Sites VisitedSites20 Total (5 Light / 15 Dark) Single Site Goal Masskg23 Payload Subsystem Total Masskg324 SMD Masskg260 ESMD Masskg64 Payload Percentage of Total Mass%32.48 Power Subsystem TypeN/ASolar Cells & Lithium Ion Batteries Total Power Masskg Total Power RequiredWe200 We Number of Solar ArraysN/A1 Solar Array Mass / Solar Arraykg Solar Array Area / Solar Arraym2m Number of BatteriesN/A24 Batter Mass / Batterykg9.824

Critical Parameters of LOW ParameterUnitsNotes Structure Subsystem Total Masskg140 Maximum G LoadG5 Thermal Subsystem Temperature of Rover (Cold Case)K160 – 250 Temperature of Rover (Hot Case)K226 – 277 Total MassKg40 Passive / Active SystemN/AActive GN&C Subsystem Total Masskg Accuracy%95 Power RequiredW Communication Subsystem Total Masskg5 TypeN/AS – Band Transmitter BandwithMHz1700 – 2300 Power RequiredW30 Data RateBits Per Minute (bpm)

Critical Parameters of LOW ParameterUnitsNotes Mobility System Range of Velocitym / hr5 – 6 Total Masskg100

Power Allocation Table Average LoadLandingInitializingDrivingScienceCommunicationsSleepEOM & SRV Thermal GN&C Power10 1 C&DH Communications Propulsion Mobility Mechanisms Payload SRV Total

CDD Requirements Atlas V-401 EPF with landed mass of 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 Subsystem Top 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 2.217m 2 Silicon Mass of kg Solar Power Required 455 W e Power Storage –Lithium Ion Batteries 24 Batteries Mass of kg per battery 45 kW e -hr Capacity During the Dark –8 days of power at 200 W e average –6 days of sleep phase at 20 W e

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 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 Phase Battery Power GN&C Landing Legs Initializing Initializing Phase Com. Sys. “OK” Solar Power Science Box Cameras on 1 st Science Site Science Phase Experiments Com. Phase Data Transfer 1 day 1 st day

Light Sites Travel to Site Traveling Phase Mobility Solar Power Charge Batteries Com. Phase Data Transfer Repeat for Light Science Site Science Phase Experiments Com. Phase Data Transfer 4 days 1 day

Dark Sites Travel to Dark Traveling Phase Mobility Battery Power Science Site Science Phase Experiments Com. Phase Data Transfer Travel to Light Traveling Phase Mobility Battery Power Com. Phase Data Transfer Await Sunlight Sleep Phase Minimal Thermal And power Repeat for Dark 4 days 1 day 3 days

Shackleton Crater Travel to EoM Traveling Phase Mobility Battery Power Solar Power Shackleton Science Phase Experiments Com. Phase Data Transfer EoM & SRV Phase Launch SRV 3 Weeks

Radiator Temperature

Radiator Temperature vs. Area

Cold Case

Hot Case

Expected Temperature Ranges