Project X pedition Spacecraft Senior Design – Spring

Motivation: Lunar Payload Delivery Resupply Lunar BaseSmall Payload

Project X pedition Requirements Land on the Moon Move 500 meters Transmit HD pictures and video to Earth Survive the Lunar Night Minimize cost with 90% success Project Xpedition

Payloads 100 g 10 kg 1700 kg

Mission Phases Earth Launch Lunar Transfer Lunar Descent Locomotion 500m

Earth Launch Dnepr ft 160 ft Falcon ft

Earth Launch Site: Baikonur Cosmodrome, Kazakhstan Cost: $5M 250 Mile Parking Orbit

Lunar Lander Orbital Transfer Vehicle 880 lbs 8’

Solar Arrays unfold Internal View Hall Thruster produces 80 mN of thrust

Power Communication S-Band Antenna 2 Solar Arrays Lithium-Ion Battery Attitude Chemical Thrusters Sun Sensor Star Sensor Reaction Wheels Lunar Transfer

16 mile parking orbit 2 hour orbital period Lander is self sufficient 350 lb Lander mass Half of mass is propellant

Space Balls Housing Communication Antenna and Motor Solar Panel Attitude Control Thrusters Radiator Attitude Sensors CPU H 2 O 2 Tank Helium Tank Radial Flow Hybrid Engine Camera

Surveyor 3 Apollo miles Landing Site: Mare Cognitum

Final Descent Attitude: 12 Control Thrusters Translation: Radial Flow Hybrid Engine Mission Requirements Land on Moon Move Payload 500 m Survive Lunar Night

Lexan Shell Camera CPU Dust Removal Vibration Motor Battery 100g Payload Main Axel and Motor Housing Communications Transceiver

Mission Requirements 1.Land on Moon 2.Move 500m 3.Take Picture 4.Survive Night Taking Photo of Lander Removing Dust All Systems Are GO! Avoiding Obstacle Cruise Speed: 3.2 mph Minimum Turning Radius: 2.5 in -280 °F

10 kg Lunar Lander 230 lbs Lander 270 lbs Propellant 500 lbs Total Mission Requirements 1.Land on Moon 2.Move 500m 3.Take Picture 4. Survive Night Hybrid Engine Thrust: 45 lbs Burn Time: 135 sec 10 kg Payload

500m Record Video Mission Requirements 1.Land on Moon 2.Move 500m 3.Take Picture 4.Survive Night

Completed lunar descent Full stop Begin locomotion Attitude Thrusters 16 ft 6 ft 300 ft Main Engine Avg. Thrust: 230 lbs Burn time: 60 s Large Payload

Mission Requirements: 1.Move 500 meters 2.Land on moon 3.Resupply base

$27M Cost - $22M Prize = $5M Net Mission Cost Mission Cost $27 Million 72% Success $30 Million 72% Success $223 Million 92% Success Cost Per Kilogram $271Million $3 Million $130k

Payload Delivery: 1. Most economical payload: 2 tons 2. Electric Propulsion for Lunar transfer 3. Soft land on Lunar surface Google Lunar X PRIZE: 1.Several viable locomotion methods 2.Potential to open commercial market 3.$27M mission accomplished for $5M Project X pedition Results

Question & Answer Project X pedition

Backup Slide Listing Propulsion Brad Appel Thaddaeus Halsmer Ryan Lehto Saad Tanvir Attitude Brian Erson Kris Ezra Christine Troy Brittany Waletzko Power Tony Cofer Adham Fahkry Jeff Knowlton Ian Meginnis Structures & Thermal Kelly Leffel Caitlyn McKay Ryan Nelson Communications Mike Christopher John Dixon Trent Muller Mission Operations John Aitchison Cory Alban Levi Brown Andrew Damon Alex Whiteman Solomon Westerman

Backup Slides Saad Tanvir Return to Listing

Propulsion System Mass Finals 100 g Payload case (Ball) Propellant mass = 78.2 kg Propulsion System Inert mass = 29.9 kg Total Prop System Mass = kg Arbitrary Payload case (Falcon 9) Propellant mass = kg Propulsion System Inert mass = 227 kg Total Prop System Mass = kg 10 kg Payload case (Hopper) Propellant mass = kg Propulsion System Inert mass = 45.4 kg Total Prop System Mass = kg 2 Saad Tanvir Propulsion Group Return to Listing

100 g – Hybrid Propulsion System Mass Breakdown 3 Return to Listing

4 10 kg – Hybrid Propulsion System Mass Breakdown Return to Listing

5 Large payload – Hybrid Propulsion System Mass Breakdown Return to Listing

Propellant Tank Specifications 6 Return to Listing

Pressurant Tank Specifications 7 Return to Listing

8 Hydrogen Peroxide Tanks - Thermodynamic Analysis Assumptions: Tank operating Temperature = 283 K (50 F) Surrounding Temperature = 2.73 K Power Required ~ 35 W ΔT = K Q: Rate of Heat transfer [W] A: Area of Cross section of the tank [ m 2 ] k: Thermal Conductivity [0.044 W/mK] ΔT: Temperature Difference [K] t: Thickness of the blanket [200 mm] Return to Listing

Lunar Descent – Thermodynamic Analysis on Prop System 9 Temperature Drop < 5 K No power required to heat the propulsion system during Lunar Descent Return to Listing

Propellant Tank – Operating Pressure P chamber = 2.07 MPa ∆P dynamic = ½ v 2 ~ MPa ∆P feed (Upper bound) ~ 0.05 MPa ∆P cool ~ 0.15pc = 0.31 MPa ∆P injector ~ 0.3pc = 0.62 Mpa P tank ~ 3.07 MPa 10 Return to Listing

Lunar Transfer: Chemical Alternative Significant mass savings using the Electric Propulsion system 11 Return to Listing

Backup Slides Christine Troy Return to Listing

Lander Attitude Control 12 General Kinetics H 2 O 2 thrusters Lander Side view Lander Top view Return to Listing

Attitude Prop Mass Estimate Based on Rauschenbakh, Ovchinnikov, and McKenna-Lawlor θ. θ +θ1+θ1 -θ1-θ1 No External Torque θ +θ1+θ1 -θ1-θ1 “Large” External Torque θ. M b = external moment applied g = gravitational acceleration Isp = specific impulse of thrusters L = distance from thruster to vehicle center of mass Return to Listing

Spinning Lander Attitude Control  Propellant and thrusters still needed for spin up and axis reorientation –Estimate ~2.2 kg propellant savings for 100g/10kg cases  Additional mass: spinning landing gear, propulsion system redesigns, additional attitude sensing devices  Increased complexity: Liquid propellant feed while spinning, landing while spinning, reorientation of axis Return to Listing

Compressed Gas Spring Energy Storage Some or all travel could be obtained from bouncing using stored descent energy Compressed gas not recommended – highly temperature sensitive, limited velocity and acceleration inputs – Commercial gas springs limited to approx. -23° to 82° Lunar surface temperature -153° to 107° C Return to Listing

Backup Slides Brittany Waletzko Return to Listing

System Masses Mass100g10 kgLarge Injected Mass to Low Earth Orbit (kg) Injected Mass to Low Lunar Orbit (kg) Mass on Lunar Surface (kg) Payload Delivered to Lunar Surface 100g10kg1743kg Systems Overview 100g Payload 10kg Payload Large Payload Return to Listing

Mission Timelines (Backup) Elapsed Time (ddd:hh:mm) EventVehicle -365:00:00LaunchLaunch Vehicle/OTV 000:00:00Arrive in LLOOTV 000:00:03In lower orbitLander 000:00:04Rotate and LandLander 000:00:04Systems checkSpace Ball 000:00:05Deployment from LanderSpace Ball 000:00:06OrientationSpace Ball 000:00:06Travel 500mSpace Ball 000:00:14 Braking maneuver, dust removal Space Ball 000:00:15 Take picture of Lander, Begin transmission to Lander Space Ball 000:00:23End photo transmissionSpace Ball 000:00:23 Transmit arrival Mooncast (near real-time video, photos, HD video, XPF set asides, data uplink set) to Earth Lander 001:33:56 Transmit Mission Complete Mooncast (near real time video, photos, HD video) Lander 002:08:04 Finished transmitting, prepare for night Lander 009:00:00Standby for lunar nightLander 025:00:00Power up after nightLander 026:00:00Transmit telemetry and photoLander 026:00:14Mission Complete Elapsed Time (ddd:hh:mm) Event -365:00:00Launch 0:00:00Lunar Lander reaches LLO and separates from OTV 0:00:04Lands on lunar surface and starts video taping 0:00:12Finishes taping and begins transmission of video 0:03:44Completes video transmission and takes panoramic pictures 0:03:45Finishes panoramic pictures and begins transmission of pictures 0:03:59Completes picture transmission and begins hop for locomotion 0:04:01Locomotion phase complete and begins HD video taping 0:12:01Begins transmission of HD video and takes panoramic pictures 2:06:24Ends transmission of HD video and begins transmission of pictures 2:06:36Ends transmission of pictures and shuts down for lunar night 15:23:24Turns on and sends signal after lunar night. Elapsed Time given in days, hours, and minutes 100g Payload Mission Timeline10kg Payload Mission Timeline 100g and 10kg Payload Return to Listing

Mission Timelines—cont. (Backup) Elapsed Time (ddd:hh:mm) Event -365:00:00Launch 000:00:00Arrive in Low Lunar Orbit Transfer to Lunar Descent Transfer Orbit Begin Final Lunar Descent burn Come to rest 100 m above surface/begin hover locomotion Touch down on lunar surface Large Payload Mission Timeline Elapsed Time given in days, hours, and minutes Large Payload Return to Listing

Trajectory Correction (backup) 100g Payload Correction Maneuver Configuration ParameterValue I sp (s)1952 m o (kg)436.0 Propellant for Correction (kg)1.1 Thrust per Engine (mN)75 Time for ΔV (hr) kg Payload Correction Maneuver Configuration ParameterValue I sp (s)1964 m o (kg)585.6 Propellant for Correction (kg)1.5 Thrust per Engine (mN)75 Time for ΔV (hr)92.6 Large Payload Correction Maneuver Configuration ParameterValue I sp (s)2250 m o (kg)9953 Propellant for Correction (kg)22.5 Thrust per Engine (mN) (x4 engines) 424 Time for ΔV (hr)54.2 T = instantaneous thrust (assumed constant over interval) m = instantaneous mass (assumed constant over interval) Return to Listing

Thruster Locations and Thrust Direction Vectors Return to Listing

Hydrazine and Hydrogen Peroxide Thrusters Return to Listing

Environmental Forces Codes Return to Listing

Output (in Newtons, Kilograms) “Environmental” : Felec = e-005 Fref = e-022 Ftherm = e-022 Fscrad = e-006 Fswind = e-009 Fmag = e-013 Fexp = e-007 Ftotal = e-005  “Environmentalpropmass” :  mm_cyl_month =   mm_cyl_5000 =   mm_cyl_50000 =   mm_cube_month =   mm_cube_5000 =  e-004  mm_cube_50000 =  Return to Listing

Backup Slides Ian Meginnis Return to Listing

OTV Power Subsystems PPU (Electric Propulsion) PCDU Battery (LL) 100V DC-DC Converters Individual OTV Components Acronym Definitions: PCDU - Power Conditioning and Distribution Unit PPU - Power Processing Unit LL - Lunar Lander DC - Direct Current Solar Array <200V Solar Array Note: Not to scale Return to Listing

GroupPower (Watts) Propulsion1529 CommunicationSee “Lunar Lander” Attitude101.4 Power120 Lunar Lander (during Lunar Transfer) 105 TOTAL g Payload Case Power Budget GroupPower (Watts) Propulsion2029 CommunicationSee “Lunar Lander” Attitude145.4 Power120 Lunar Lander (during Lunar Transfer) 105 TOTAL kg Payload Case Power Budget GroupPower (Watts) Propulsion38773 CommunicationSee “Lunar Lander” Attitude305.4 Power Lunar Lander (during Lunar Transfer) 105 TOTAL42960 Large Payload Case Power Budget Return to Listing

Power Distribution: 100g Payload OTVPower Distribution: 10kg Payload OTV Power Distribution: Large Payload OTV Return to Listing

Payload SizeComponentVariableValue 100g Solar Arrays (2 circular arrays) Mass13.06kg Deployed Area6.54m 2 Cost$1.96 Million Battery Mass12.17kg Cost$22,000 10kg Solar Arrays (2 circular arrays) Mass16.89kg Deployed Area8.45m 2 Cost$2.53 Million Battery Mass15.9kg Cost$28,600 Large Solar Arrays (2 rectangular arrays) Mass286.4kg Deployed Area143.2m 2 Cost$42.96 Million Battery Mass271.7kg Cost$488,400 OTV Power Dimensions Return to Listing

Note: Not to scale Acronym Definitions: PCDU - Power Conditioning and Distribution Unit PPU - Power Processing Unit Batt - Battery DC - Direct Current PPU PCDU Batt DC/DC Converter Aluminum Heat Pipes with Ammonia Aluminum Mount 2 Radiators Electronics Board Thermal Control Hall Thruster Thermal Control Hall Thruster Radiating Heat Shroud Radiating Heat Shroud (Exhaust) OTV Thermal Control (all payloads) Return to Listing

OTV Electronics Thermal Control Payload SizeComponentMass (kg) 100g Radiators1.2 AmmoniaNegligible Heat Pipes2.2 TOTALS3.4 10kg Radiators1.6 AmmoniaNegligible Heat Pipes2.5 TOTALS4.1 Large Radiators23.4 AmmoniaNegligible Heat Pipes15.3 TOTALS38.7 Return to Listing

Note: Not to scale At least 1 of the OTV’s set of radiators will not be exposed to sun’s rays at any point during the trajectory Each radiator, alone, can provide thermal control for OTV electronics Earth Sun Moon Single, Simplified Orbit of OTV (Large Payload) Return to Listing

Backup Slides Ryan Nelson Return to Listing

Basic Frame Design Drivers in frame mass – Total Lunar Lander (LL) mass at lunar touchdown – Volume of LL Shape: Conic Frustum – Stores all Lunar Lander subsystems while minimizing volume All frame components hollow – Small leg diameter allows for storage within side supports prior to lunar touchdown Schematic of frame listing components Return to Listing

Final Mass and Volumes Lunar LanderVolumeHeightTop DiameterBottom DiameterMass 100g1.05 m m 1.3 m11.44 kg 10kg 1.15 m m1.0 m1.3 m19.97 kg Large m m2.4 m 3.6 m kg Return to Listing

Frame Design Thickness of all frame components varies – First mode of failure (factor of safety = 1.5) – Payload case Cross sectional shape is circular or rectangular for all components 0.5 mm magnesium skin place around Lunar Lander frame – Micrometeorite protection – Thermal protection Return to Listing

Floor Supports Support a majority of landing loads Thickness altered until load is supported Hollow Rectangular Cross section – Moment of Inertia – Bending Stress acting on beam Return to Listing

Side Supports/Legs First mode of failure is buckling – K=0.5 for side supports (both ends fixed) – K=2.0 for legs (one end is free to move) – Hollowing the rod decreases moment of Inertia and critical load F cr = Return to Listing

Side Supports/Legs Compression failure occurs after buckling for both side supports and legs – Despite small cross sectional area – Compression failure Top, Bottom, and Engine Support rings all designed to have same cross sectional dimensions Return to Listing

Cross Sectional Dimensions 100g Payload Case Structural ComponentCross sectional shapeOuter DimensionsThickness Outer RingCircular10 cm Diameter4mm Engine Support RingCircular10 cm Diameter4mm Rectangular Floor SupportsRectangular10 cm Height, 6 cm Width6mm Side SupportsCircular10 cm Diameter2mm Top RingCircular10 cm Diameter4mm LegsCircular6 cm Diameter3mm 10kg Payload Case Structural ComponentCross sectional shapeOuter DimensionsThickness Outer RingCircular10 cm Diameter5mm Engine support RingCircular10 cm Diameter5mm Rectangular Floor supportsRectangular10 cm Height, 6 cm Width7mm Side supportsCircular10 cm Diameter2mm Top RingCircular10 cm Diameter5mm LegsCircular5 cm Diameter3mm Return to Listing

Placement of Hop Engines Return to Listing

Backup Slides Brian Erson Return to Listing

Backup Slide 1 Calculation of Thrust Misalignment Torque Estimate of Thrust at 1 kW~100 mN Estimate of Thrust misalignment~ 0.05 m Conservative Max Misalignment Torque ~ 5 mNm Calculation of Drift Error Tracking error from Reaction wheel spec sheet<1rpm Operating speed3000 rpm Max Wheel Torque 12 mNm Drift Error = 1/3000 * 12 = mNm Return to Listing

Backup Slide 2 Calculation of Mass Requirement (Ref, Smart-1 Lunar Probe) Reaction Wheel assembly~ 12 kg Sun Sensors~ 4 kg Angular rate sensors~ 0.3 kg Star Tracker~ 3 kg Total~ 19.3 kg ACS/Launch mass: 19.3/380 =.05 = 5% Conservative estimate of IMTLI: 700 kg * 5% Conservative estimate of ACS mass < 35 kg Return to Listing

Backup Slide 3 Calculation of Pointing Accuracy Max pointing error of SMART-1:60 arcminutes 1 arcminute = 1/60*deg 1 deg =.017 rad Return to Listing

Backup Slide 4 Attitude Control Mass Calculations: 3-axis Sun Sensors0.7 Star Sensors6.4 Reaction Wheels6 H2O2 Thrusters1 Propellant(H2O2)*26 Total:40.1 kg Spin Conical Scanner6 Doppler Device1 H2O2 Thrusters1 Propellant(H202)**19.1 Total:27.1 *Prop Mass Includes Lunar Descent Return to Listing

Backup Slide 5 **H2O2 De-Saturation(DS) Mass Calculation: DS of reaction wheels: Estimate of DS maneuvers/day:6 Reaction wheel max torque:0.03 N-m H2O2 Thrust:9.5 N/kg Max Mission Length:365 days Total Mission DS H2O2 mass: (365)(6)(1/9.5)(0.03) = 6.9 kg Return to Listing

Backup Slide 6 Other Mass Calculations: Power: Masses based on posted Power Group Data Mass of battery without solar cells based on assumption of >1.2 kW needed to power OTV Communication: 3 kg mass based on posted Com Group Data Thermal: 3 axis mass based on posted data Spin stabilized Thermal Protection: Assume 15 rpm Mass = [(1/15(17.1))+4*] =~ 5 kg *Estimated mass of standard thermal protection Return to Listing

Backup Slide 7 Cost tradeoff: Xe thruster system cost:$100,000 Current Earth to LEO cost/kg:$4400 Economical mass savings 22.7 kg Xe system mass savings < 5.0 kg Xe system Earth to LEO cost:~$20,000/kg Note: Unless Xe system saves upward of 22.7 kg,or the cost decreases, it is not economically feasible to install the system. Further analysis will be done to improve mass and dollar cost numbers of both systems. Return to Listing

Backup Slide 8 Xe DS propellant calculation Total DS force needed: Max torque of Rxn Wheel:0.03Nm DS per day:6 Mission length:365 days Moment arm1.0m Total:65.7 N Total number of N4380thrusts Marotta Cold Xe gas thruster Specs: Mass:0.075kg Isp:68sec Thrust:0.015N Time per thrust:0.04sec Mass Flow Calculation: Isp = Force/massflow * gravity Mass flow:0.0022kg/sec Total Mass: kg/sec * 0.04 sec/thrust * 4380 thrusts = kg Return to Listing

Thruster Analysis 100g10kgArbitrary Added Inert mass (kg) Added Volume (m^3)4.3x x x10 -4 Cost savings($) Consultation with Purdue Hybrid(H202) Rocket Team led to development of an alternate OTV attitude control system System consists of 4 small H202 tanks enclosed within OTV Each system is independent All payload cases can be developed in-house for a fraction of purchase cost Backup Slide 9 Return to Listing

Reaction Wheel Update Payload DeviceManufacturerMass (kg)Size (cm)Power Required peak (W)Max Torque (mNm) 100g VF MR 4.0 (4)Valley Forge Composites2.6 (each)20 x10 (each)76 (total)20 (each) 10kg VF MR 10.0 (4)Valley Forge Composites5.0 (each)25 x15 (each)120 (total)30 (each) Arbitrary VF MR 19.6 (4)Valley Forge Composites10.5 (each)39 x17 (each)280 (total)260 (each) Each Reaction Wheel had to be upgraded within each payload to account for increases in system mass Relevant changes to note: 100g10kgArbitrary Mass Increase (kg) Power Increase (W) Backup Slide 9 Return to Listing

Backup Slide 10 Cost Savings Calculation: 100g General Kinetics Cost for 4 – 1N thrusters:$12,000 In-house Manufacturing cost:$5,000 Cost Savings:$7,000 10kg General Kinetics Cost for 4 – 1N thrusters:$12,000 In-house Manufacturing cost:$6,000 Cost Savings:$6,000 Arbitrary General Kinetics Cost for 4 – 13N thrusters:$48,000 In-house Manufacturing cost:$10,000 Cost Savings:$38,000 Return to Listing

Backup Slide 11 Inert Mass Calculations Density of H202:1.11 kg/L Mass of aluminum tank per.001 m^3:3.68kg Kg(prop) = massflow*(sec/thrust)*thrusts Kg(tank) = (3.68/.001)*volumeH g 4 – kg H202 Tanks0.064kg 4 – 0.02N H202 Thrusters0.36kg Feed Lines, Valves1.5kg Total Inert Mass1.924kg 10kg 4 – 0.315kg H202 Tanks1.16kg 4 – 0.03N H202 Thrusters0.36kg Feed Lines, Valves1.5kg Total Inert Mass3.02kg Arbitrary 4 – 2.65kg H202 Tanks8.8kg 4 – 0.26N H202 Thrusters0.36kg Feed Lines, Valves1.5kg Total Inert Mass13.84kg Return to Listing

Backup Slides Caitlyn McKay Return to Listing

Deployment Linear Shaped Charge SystemMass (kg) / Space Ball Charge0.580 Foam0.040 Total Return to Listing

Accordion Landing Return to Listing

Accordion Landing Return to Listing

Impulse Momentum Return to Listing

Kamikaze Rover Solar Panels0 kg Batteries0.422kg Power (extra)1.92kg Communications1.51kg Drive System0.298kg Structure (frame)0.40kg Space Blankets0.58kg Wheels0.58kg Cooling System0kg System Mass5.7074kg Ballast Mass10kg Total kg Length0.23m Width0.21m Height0.21m * Life of 13 minutes. Return to Listing

Rover Deployment Linear Shaped Charge System to lower Rover from Lander to surface. ItemLinear Shaped Charge SOLIMIDE Foam Steel Cable PlatformMotorSupport Beams Total Mass (kg) Return to Listing

Backup Slides Trent Muller Return to Listing

1.Mt. Pleasant Radio Observatory. Hobart, Tasmania, Australia. A 26 meter dish. 2.Hartebeesthoek Radio Astronomy Observatory (HRAO). Johannesburg, South Africa. A 26 meter dish. 3.Pisgah Astronomical Research Institute (PARI). Rosman, North Carolina. USA. One of the 26 meter dishes. 4.James Clark Maxwell Telescope. Mauna Kea Observatory, Hawaii, USA. A 15 meter dish. Return to Listing

Ground StationsAltitude (km) of Non-Tracking Zone Ground Station Latitude ( o )Longitude ( o ) S E S27.68 E N82.87 W N W Communications Coverage Return to Listing

EquipmentModelManufacturerMass (kg)Power Usage (W)Price (2009 $) Lander-Earth Antenna (2) Patch AntennaSSTL ,000 Lander-Earth Receiver RX-200SSpaceQuest ,000 Lander-Earth Transmitter TX-2400SpaceQuest ,000 Lander-Rover Antenna ANT-100SpaceQuest Lander-Rover Transceiver TR-400SpaceQuest ,000 Computer BoardRAD6000BAE ,000 Video CameraHF10Canon ,000 Antenna Pivot (2) Totals ,668 Communications Equipment Onboard Lander for 100 g Payload Return to Listing

Communications Equipment Onboard Lander for 10kg and Large Payload EquipmentModelManufacturerMass (kg)Power Usage (W)Price (2009 $) Lander-Earth Antenna (2) Patch AntennaSSTL ,000 Lander-Earth Receiver RX-200SSpaceQuest ,000 Lander-Earth Transmitter TX-2400SpaceQuest ,000 Computer BoardRAD6000BAE ,000 Video CameraHF10Canon ,000 Antenna Pivot (2) Totals ,167 Return to Listing

Backup Slides Tony Cofer Return to Listing

Hydrazine Heater for 100g and 10kg Payloads Return to Listing

Nocturnal Power Controller Save 11.5 kg of batteries Controllable Size 2”X2”X1/2” Weight~20 g Power diss. 0.1mW Requires 0.23 g battery for 14 days Controller Interface Solar Source Comparator With Hysteresis Actuator Command Computer Return to Listing

Backup Slides Mike Christopher Return to Listing

Mooncast Schedule Lunar Arrival Mooncast ItemLink DirectionSize [MB]Transmission Time [hr] PhotosDown50.24 XPF Set AsidesDown Data Uplink SetUp Data Uplink SetDown Totals35 MB1.65 hrs Locomotion Mooncast ItemLink DirectionSize [MB]Transmission Time [hr] 8 min Near Real Time VideoDown min High Definition VideoDown PhotosDown50.24 Totals980 MB46.14 hrs Survival Mooncast (BONUS PRIZE) ItemLink DirectionSize [MB]Transmission Time [hr] 8 min Near Real Time VideoDown PhotosDown50.24 Totals80 MB3.77 hrs Michael Christopher – Backup Slide Return to Listing

Patch Antenna and Pivot System Advantages  Redundancy (2 pivots and antennae and 2 motors on each pivot.  Reduces the need for more antennae on the Orbital Transfer Vehicle (OTV)  Low cost pivot: $83.50  Low mass pivot: 0.2 kg System Mounted on OTV Michael Christopher – Backup Slide Return to Listing

Antenna MountBase Plate Stepper Gear Motor Patch Antenna and Pivot System Michael Christopher – Backup Slide Return to Listing

Mass: 2*( kg/motor) + 0.1kg = kg Power Consumption: Watts Cost: 2*($16.75 /motor) + ~$50 Al = $83.50 Patch Antenna and Pivot System Michael Christopher – Backup Slide Return to Listing

Michael Christopher – Backup Slide Patch Antenna and Pivot System Return to Listing

Backup Slides John Aitchison Return to Listing

Lunar Descent Overview Note: Not to Scale Lunar Parking Orbit Lunar Descent Transfer Orbit Final Descent Moon Return to Listing

Final Descent Overview Return to Listing

Final Descent Trajectory Return to Listing

100 g Payload Descent Overview Return to Listing

10 kg Payload Descent Overview Return to Listing

1743 kg Payload Descent Overview Return to Listing

Descent Validity Check ∆V ~ 2,000 m/s to move from LPO to zero velocity on lunar surface I sp = 320 s g 0 = 9.8 m/s 2 M i = Total Lander Mass in LPO = 157 kg. M f = 83 kg Propellant Used = M i – M f = 74 kg Return to Listing

Equations of Motion Return to Listing

Altitude & Range Return to Listing

Surface Clearance: Worst Case Scenario Return to Listing

Lander Mass vs. Time Return to Listing

Lunar Descent Transfer Orbit Return to Listing

Sample Descent Code Output Return to Listing

Backup Slides John Dixon

Thermal Considerations Assumptions: – Solar Panels Reflect Unused Solar Energy Completely – Thermal Blanket keeps Energy Transfer through body to 0 J/s Above includes MLI comprised of Kapton (or Teflon) / Silver Lined Reflective Surface, Kapton Insulation (with scrim separation) – Thermal Heat Sinks radiate to Coldest Possible Surface – Steady State Conduction Return to Listing

Insulation/Heat Sink Copper Heat Emission – q/t = J/s (from emissivity of Copper) – Cu mass = 6.08 kg One heat vane traveling to each side of the CPU Operating Temp Multi-Layer Insulation (MLI) – Insulation mass= kg Total Thermal Control Mass: 6.98 kg Return to Listing

Copper Sink Properties Copper Slab – 0.03m thick X 0.065m wide X 0.08m long – Volume: m^3 Copper Vein – 0.02m height X 0.065m wide X [0.001:0.372]m thick – Volume(max distance) = m^3 Return to Listing

System Description N 2 1 atm inside Toy Ball Enclosure – Mass of N 2 gas = kg – Temp of N 2 gas = 0 o C (273.15K) Total Heat Dissipation – Z-93 White Paint Coating (α = 0.17) Q sun = J Q electronics = 1060 J Q total = J Total Energy Rate Into System = W Return to Listing

Thermal Transport Over Time Steady State Equilibrium Occurs at ~50 seconds Total Temperature Rise Over 8 min = 0.7K Return to Listing

Backup Slide 1 Return to Listing

Backup Slide 2 Return to Listing

Lander to Earth Transmission Distance from 200 km Parking Orbit to 440,000 km of Moon at Apogee Transmit Satellite: m Receiver Satellite: DSN 26 m Minimum Power: 33 W Frequency: 2.2 GHz (S-Band Range) Data Rate: 51.2 kbps Return to Listing

Rover to Lander Transmission Distance from 0 m Lander to 500 m Maximum Travel Transmit Satellite: m Receiver Satellite: Lander 0.2 m Minimum Power: Open Condition Frequency: 2.2 GHz (S-Band) Data Rate: 51.2 kbps Return to Listing

Beamwidth Optimization (Backup) Return to Listing

Backup Slides Jeff Knowlton Return to Listing

Overview 1 Minute Deploy 2 Status Relays 8 Minutes Travel 1 minute Prep/ Photograph 8 Minutes Transmitting Space Ball Power Return to Listing

Ball Power System Battery (using three) Lithium Manganese Dioxide Coin (CR2330 ) 3 volts.26ampere-hr Cylinder Dimensions 23mm diameter 3mm height kg each 5% loss per month(self discharge 1 year) Total 2.34watt-hr at Liftoff 45.96% loss over 1 year 1.26 Watt-hrs after 1 year 0.112kg including housing Return to Listing

Backup Slides Thaddaeus Halsmer Return to Listing

(2) (3) (4) (1) Table 1 Engine performance parameters Engine No.Payload case/DescriptionF_max/min [N]tb [s] 110 kg/hop engine 2x192 (avg.) g/main engine1100/ kg/main engine1650/ Arbitrary/main engine27000/ Stick is 6.5 feet high, same as a standard doorway Lunar Lander Propulsion – Engine Specifications Return to Listing

SV01 SV02 High Pressure Helium Tank HV01 REG CK01 CK02 MOV F01 H 2 O 2 Tank HV02 RV01 Lunar Lander Propulsion –fluid system diagrams SV01 SV02 High Pressure Helium Tank HV01 REG CK01 CK02 MOV F01 H 2 O 2 Tank HV02 RV01 SV04 SV03SV05 100g and Large payload cases10kg payload case Return to Listing

Figure X: Propellant mass vs. I sp trade Lunar Lander Propulsion - Propellant/Propulsion system selection Selection Criteria: 1.Thrust a.min/max b.throttling 2.Dimensions a.Short and fat 3.Mass – minimize 4.Propellant storability 5.Purchase/development costs 6.High Reliability Return to Listing

As area ratio, ε, increases M nozzle increases, but I sp increases also As I sp increases M prop decreases for a given thrust and burn time Wrote Matlab script that used Matlab CEA interface to compute multiple I sp ’s for different area ratio’s and the corresponding M prop and M nozzle for a given thrust, and burn time Results: Area ratio for minimum mass occurred at ~150, however this nozzle would be very large and little is gained above ~100 Lunar Lander Propulsion - Nozzle area ratio and mass optimization  Used CEA to compute I sp for given nozzle area ratio All other inputs constant  Empirical nozzle mass equation Return to Listing

Fuel grain dimension definitions Lunar Lander Propulsion – I sp analysis approach Return to Listing

Lunar Lander Propulsion – fuel grain and chamber sizing approach 1.Choose a.Empirical value for initial fuel regression rate b.Initial O/F ratio for optimum I sp c.Initial propellant mass flow rate Compute required burn surface area 2.Dimensions of fuel grains a.Diameter is derived from burn surface area found from values in step #1 and chosen fuel grain geometry b.Thickness is function of burn time and regression rate 3.Compute Chamber dimensions a. Chamber dimensions approximated from fuel grain size and additional room for insulating materials Return to Listing

Backup Slides Alex Whiteman Return to Listing

LiftoffTouchdown 10kg Hop Trajectory Trajectory Timeline First, throttle up and then throttle down engine while pitching Lander in clockwise direction. Next, Lander remains at constant pitch angle and altitude while thrusting in direction opposite of hop Finally, Lander pitches in counter-clockwise direction in order to land in a vertical orientation. Return to Listing

Large Payload Hover Trajectory Trajectory Timeline First, Attitude control system moves Lander horizontally while slowly descending. Next, Attitude control system thrusts in opposite direction to cancel horizontal velocity. Main engine fires to cancel vertical velocity Return to Listing

Hopper Trajectory Results Backup Slide 1 Return to Listing

Hover Trajectory Results Backup Slide 2 Return to Listing

EOM’s r = distance of the Lander from the center of the moon θ = angular displacement along the surface of the moon measured from the start of the trajectory μ = gravitational parameter of the moon equal to km 3 /s 2 T radial = thrust in the radial (r) direction T theta = thrust in the angular (θ) direction m = mass of Lander Return to Listing

Hover Trajectory Assumptions and Constraints Initially, Lander comes to complete stop 100m above lunar surface Lander remains in upright position throughout trajectory Lander touches down with near zero horizontal and vertical velocity Main lunar descent engine responsible for all vertical movement Attitude control system responsible for all horizontal movement Lander must cover 500m distance in greater than 60 seconds Horizontal velocity limited by maximum thrust provided by attitude control system Return to Listing

range (m) altitude (m) Moon Hop Trajectory 10kg Hop Trajectory LiftoffTouchdown Return to Listing

Hop Trajectory Assumptions and Constraints 2-D trajectory in plane normal to Moon’s surface Instantaneous throttling of hybrid engine Lander takes off and touches down with near zero vertical and horizontal velocity and upright orientation Rotation rate of Lander limited by torque provided by attitude control system Return to Listing

Hop Trajectory Design In order to maintain 90% chance of success, cannot relight main lunar descent engine to perform hop. Instead use pair of redundant thrusters to perform hop. Unusual trajectory shape due to thruster configuration: one thruster firing at 32° from vertical. Must offset thrust direction by having Lander velocity in opposite direction to ensure no horizontal velocity upon landing. With out this trajectory shape, Lander would crash and/or land on its side. This trajectory adds only 2.5kg of propellant compared to a trajectory using a vertical thruster. Return to Listing

Backup Slides Cory Alban Return to Listing

Completion of Mission Requirements [Cory Alban] [Mission Ops] [Locomotion] StepTime (min)Tasks to be completed 10 Space Ball performs a system diagnosis. 21 Deployment from Lander. 32 Direction of travel received from mission control. Space Ball orients to path of travel Accelerate to cruising speed of 1.04m/s. Travel for 8 minutes until 500m objective achieved. 511 Braking maneuver with a 90 degree orientation change to point camera toward Lander. Shake off dust if necessary. 612 Snap photo of Lander from ball and begin transmission. 720 Finish Photo Transmission. RequirementSteps to Completion Travel 500m in a controlled manner1-4 Carry 100g payload 500m1-4 Transmit Mission Complete Mooncast6-7 Return to Listing

Lunar Surface Hazard Analysis Potential HazardSolution Lunar regolith Very fine dry powder Sticks to everything  Using gradual acceleration, the space ball avoids peeling out and digging into the regolith  Vibration Motor shakes off any collected regolith Impact Craters 2cm to several meters in diameter  Choose path to avoid large craters  Built up momentum reduces chance of getting caught in a crater Debris/Rocks Debris size: m to 0.50m  Lexan shell will withstand a full speed collision  At cruising speed, momentum carries ball over small rocks and retains stability (similar to a rolling wheel) Temperature Average day temperature 107C Highest day temperature 123C  Temperatures are within tolerances for Lexan  1atm of N2 inside Lexan shell to control temperature rise within the space ball Temperatures are within thermal range for Lexan Return to Listing

Space Ball Structure Analysis [Cory Alban] [Mission Ops] [Locomotion] Bending Moment in Drive Axel Model as a thin circular rod R = 0.125m Aluminum 2024 Alloy σ= 220 MPa ρ= 2730 kg/m 3 Maximum loading conditions (8.3g) g = 8.3 * m/s 2 = 75.25m/s 2 M pay = kg Minimum required radius: 1.17*10 -8 m Torsion Stress in Drive Axel Maximum Torque, T = 0.31 Nm Minimum required radius: 1.10*10 -4 m Design radius: 0.003m Factor of Safety: 27 R M pay *g T Return to Listing

Space Ball Structure Analysis [Cory Alban] [Mission Ops] [Locomotion] Sphere Impact Analysis Assume all kinetic energy converted to impact energy Cruise Speed, v = 1.04m/s Ball Mass, m = 2.435kg Total Kinetic Energy, K = 1.317J Impact Strength of Lexan, σ = 600 – 850 J/m Minimum wall thickness: 1.55*10 -3 m Pressure Vessel Analysis Pressure, P = Pa (1atm) Radius of sphere, R = m Maximum Stress of Lexan, σ = 75 Mpa Minimum wall thickness: 8.4*10 -5 m Design wall thickness: 3.82*10 -3 m Factor of safety: 2.5 R Return to Listing

Backup Slides Adham Fakhry Return to Listing

Final Power Systems 100 gram LanderMass (kg)Dimensions (m)Cost ($) Solar Cells m 2 250,000 Batteries X X DC-DC Converters X 0.05 X ,000 PCDU (Power Conditions and Distribution unit) X X , kg LanderMass (kg)Dimensions (m)Cost ($) Solar Cells m 2 250,000 Batteries X X DC-DC Converters X 0.05 X ,500 PCDU (Power Conditions and Distribution unit) X X ,000 Return to Listing

Final Power Systems for Arbitrary Arbitrary LanderMass (kg)Dimensions (m)Cost ($) Solar Cells m 2 250,000 Batteries X X DC-DC Converters X 0.06 X ,000 PCDU (Power Conditions and Distribution unit) X X ,000 Return to Listing

Backup Slide 1: Power Available to the Lander Return to Listing

Battery Design (1) Battery is designed for meet three power goals for 100 g Lander: – Delivers 124 W for 250 seconds for operating the Lander engine – Delivers 30 W for 576 seconds of attitude – Delivers 60.4 W for 30 minutes for all communication gear Return to Listing

Battery Design (2) Battery is designed for meet three power goals for 10 kg Lander: – Delivers 150 W for 450 seconds for operating the Lander engine – Delivers 30 W for 900 seconds of attitude – Delivers 56.4 W for 30 minutes for all communication gear Return to Listing

Battery Design (3) Battery is designed for meet three power goals for Large Lander: – Delivers 275 W for 500 seconds for operating the Lander engine – Delivers 30 W for 900 seconds of attitude – Delivers 56.4 W for 30 minutes for all communication gear Return to Listing

Solar Array sizing Solar array Calculations: Dimensions of Solar cells: – Area of Lander roof = π(1/2) 2 = m 2 – Solar efficiency = 300 W/m 2 – Potential max power = W Cost of Solar Cells: – Cost of cells per watt = 1000 $/W – Cost of Cells = 235, = $235,600 – Total cost = $235, ,400 (for additional costs) = $250,000 Return to Listing

Hydrazine Tanks 100 g – 3.9 kg Hydrazine kg Tank = 4.2 kg – 0.2 m diameter tanks, V= m 3 10 kg – 4.13 kg Hydrazine kg Tank = 4.41 kg – 0.21 m diameter tanks, V= m 3 Large Payload – 42.6 kg Hydrazine kg Tank = kg – 0.43 m diameter tanks, V= m 3 Return to Listing

Battery Specifications 3.6 V, 20 Ah Lithium Ion Cell Gives 72 W-hr only need 44 W-hr Energy Density = 140 W-hr/kg Dimensions = m X m X m Cost $2000 per cell From Yardney - Lithion Return to Listing

Heats of Reaction Calculations 10 W 14 days =10W∙14 days∙24 hrs/day.60 min/s.6 secs= Joules H rxn = J/mol = J/Kg Mass of Hydrazine = 3.45 kg Return to Listing

Backup Slides Kelly Leffel Return to Listing

Schematic of Heat Transfer Return to Listing

Thermal Control Total 100 gram payload10 kg payloadArbitrary payload MLI blanket2.35 kg2.38 kg21.4 kg Heaters0.5 kg0.45 kg34.1 kg Cooling System6.72 kg6.73 kg1.03 kg TOTAL9.57 kg9.56 kg56.53 kg Return to Listing

ComponentMass (kg)Dimensions (m) MLI blanket2.35lays on equip Al plate x 0.1 m 2 Heat pipe2.65 m, Ø Radiators x x0.311 Ammonia Heaters thick 100g ComponentMass (kg) Dimensions (m) MLI blanket2.38lays on equip Al plate x 0.1 m 2 Heat pipe2.515 m, Ø Radiators x 0.38 x0.38 Ammonia Heaters thick 10kg Return to Listing

Large Payload ComponentMass (kg)Dimensions (m) MLI blanket 21.4 lays on equip Al plate x 0.1 m^2 Heat pipe m, Ø Radiators x 0.81 x 0.81 m Ammonia Heaters thick Return to Listing

MLI Blanket Lander surface, propulsion system, and space balls’ compartments (100 g) 30 layers Aluminized Mylar (0.007 g/cm^2) Effective emissivity= Q = e*(A)*sb*(Th^4-Tc^4) e = Effective emissivity = A = Surface area (changes for each lander) sb = Stefan-Boltzmann constant = 5.67 *10^-8 J/K^4.m^2.s Th = Hot temperature (temperature in the sun) = 393 K Tc = Cold temperature (temperature in the lander) = 293 K Additional 0.4 kg on the 100 g case for the ball storage box Return to Listing

Heat needed to be removed Assume 70% efficient equipment With 40 Watts required, 12 Watts of heat released Communication Equipment Heat 100g – 49 Watts 10 kg – 38 Watts Arbitrary – 282 Watts Return to Listing

Communication Equipment has a Max Temperature of 313 K, keep at 303 K as a factor of safety Keep Lander Operating Temperature around 293 K Similar Thermal Control as the OTV – Area of Plate : 0.1 m^2 – Aluminum (Al) thermal conductivity : 236 W/(m*K) – Al density: 2700 kg/m^3 – Thickness < AK(T 1 - T 2 )/q < 3.8 m (for both cases) Choose 0.5 cm ( m) – Mass of plate = density * thickness * area = 1.4 kg Aluminum Plate Return to Listing

Ammonia Latent heat of vaporization of Ammonia: 1371 kJ/kg Mass (100 g) = kW * 450 sec /(1371 kJ/kg) = 0.02 kg Mass (10 kg) = kW * 450 sec /(1371 kJ/kg) = kg Aluminum Heat Pipes (100g) Volume needed to simulate P=1 atm : m^3 Choose pipe of 5 m long m^2 cross sectional area pi*r i ^2 = m^2 : r i = m, r o = m Mass = 2700 * pi * (r o ^2 – r i ^2) * length = 3.3 kg Heat Pipes Return to Listing

Heat Pipe Continued Aluminum (10 kg) – Volume : m^3, choose length = 5 m – m^2 cross sectional area – pi*r i ^2 = m^2 : r i = m, r o = m – Mass = 2700*pi*(0.0322^ ^2)*5 = 2.7 kg Radiators – Dissipate 61 and 50 Watts – Emissivity of 0.92 for white paint – Area of the radiators: m^2(100 g) and m^2 (10kg) – Mass = 2.38kg(100 g), 1.95kg (10 kg) Return to Listing

Backup Slides Ryan Lehto Return to Listing

Space Ball Propulsion System Performance Average Velocity: 1.04 m/s (2.33 mph) Max Inclination: 14.42° Acceleration: m/s 2 Power Usage: W Turning Radius:.0625 m (2.46 in) Largest Boulder Traversable: 0.325m (12.79 in) Propulsion System Mass: kg (0.379 lbm) Return to Listing

Ball Movement Forward/Back Movement Left/Right Movement Return to Listing

Motor Data Motor Power Input (W)EfficiencyPower Nominal Output (W)Mass (g) No-load Velocity (RPM) Dia (mm)Length (mm)Cost RE 6 Ø6 mm, Precious Metal Brushes, 0.3 Watt % $58.71 (45.87 Eur) GearingRatioEfficiencyMass (g)Diameter mmLength mmCost Planetary Gearhead GP 6 A Ø6 mm 221:160% $94.55 (73.87 Eur) Sources: and Stepper Motor Holding Torque (Nm)Step AngleVoltage (V)Mass (g) No-load Velocity (RPM)Cost ARSAPE Two Phase Stepper Motor -- AM2224-R3AV ° $58.71 (45.87 Eur) Return to Listing

Alternative Propulsion Comparison Space Ball Motors: 2 (One Stepper & One Continuous D/C Motor) Additional Mass: Drive Shaft & Swing Arms kg (0.379 lbm) Largest Boulder Traversable: m (12.79 in) Rover Motors: 4 Continuous D/C Additional Mass: 4 Wheels and Motor Mounts kg (5.54 lbm) Largest Boulder Traversable: m (4.45 in) Return to Listing

Backup Slides Kris Ezra Return to Listing

Gimbaled Main Engine Alternative Gimbal Mount Specifications: 1.Approximate mass of 6 kg 2.Angular maximum motion of 20º – 3 Axis Gimbal 20º 1.06 m m Mission Length (days)365 Desaturation Maneuvers (#/day)6 Max Reation Wheel Torque (Nm)0.03 H202 Specific Thrust (N/kg)9.5 Attitude Moment Arm (m)1.1 Engine Moment Arm (m) Total mass (kg) (Attitude DS) Total mass (kg) (Engine DS) Gimbal Alternative Discarded based on Mass Cost Return to Listing

Spinning Mass Tether Alternative Rationale for Discarding Momentum Transfer Concept: The momentum transfer concept was analyzed just using work/energy relationships subject to the conditions that the Lander could not experience an acceleration greater than 10g and that the Lander would initially be traveling at an orbital speed of 1.7 km/s. Because the constraint on the system is an acceleration and the frame of the moving Lander is not inertial, the system was analyzed using work/energy but in an inertial frame. This approach has obvious limitations; however, it also should provide a more conservative analysis meaning that, if the results are unfeasible for this simplified model, the addition of a gravitational component by the moon will only make exacerbate the outcome. Shown below is a plot of the acceleration felt by the Lander versus collision/spring distance through which some force must act to slow the Lander to zero. A reasonable distance for this “collision” would be between 1 and 2 meters since a spring of this relaxed length must be carried on the OTV with a mass less than that of the Lander descent propellant. From the graph, it can be seen that, at this distance, the accelerations are on the order of 1x10 5 Earth g’s. This is four orders of magnitude higher than that sustainable by the communications equipment (10g) and is probably higher than what is able to be withstood by the molecular bonds in the vehicular structure. Additionally, to maintain an acceleration less than 10g during a deceleration from 1.7 km/s it would be necessary to have a collision distance of approximately 150 km. For these reasons among others, the momentum transfer concept is infeasible. Acceleration sustainable by Communication Equipment: 10g Required Tether Length to Match Orbital Velocity: ~50 km Additional mass cost at this Length: 325 kg (Total mass of 400 kg) Orbital Height: ~100 km Result: Weight of tether exceeds propellant mass and tether length is nearly half the orbital height. Completely infeasible. Return to Listing

Backup Slides Andrew Damon Return to Listing

X Y x y Barycenter Circular-Restricted Three-Body Problem Two coordinate frames: and fixed inertially x and y rotate with Earth-Moon system Equations of Motion (including thrust): VariableDescription xydxyd Component of OTV position in x-direction Component of OTV position in y-direction Distance from the Earth’s center to the OTV R Distance from the Moon’s center to the OTV Distance from the Earth’s center to the barycenter Gravitational parameter NMean motion of the system, normalized to 1.0 T x * T y * m* M o * Thrust in the direction of x velocity component Thrust in the direction of y velocity component Current mass of the OTV Initial mass of the OTV in Earth parking orbit Mass flow rate of the EP system *Much more accurate than patched two body model *Gravity effects of Earth and moon are always taken into account Return to Listing

Recommend: Parking Orbit of 400 km – Drag drops to less than 5% of Thrust, Within capabilities of Dnepr Launch Vehicle Assume Thrust of 110 mN Assume C D = 1.0 Analysis based on cross section area of: Solar Panels ~ 8 m 2 OTV ~ 4 m 2 Total Area ~ 12 m 2 Circular Parking Orbit Altitude (km) Drag (mN) T/D Assume Thrust of 110 mN Atmospheric Drag for Circular Parking Orbits Return to Listing

Drag Calculations F D ~ Newtons ρ ~ kg/m 3 C D ~ dimensionless v ~ m/s A ~ m 2 Backup Slides Altitude (km) Circular Velocity (km/s) Curve fit for density based on altitude: Where h is altitude in km and ρ is in ng/m 3 Return to Listing

Backup Slides Levi Brown Return to Listing

Correction Maneuver: 50 m/s Burn: Additional Propellant Requirements 100 g – 1.1 kg 10 kg – 1.5 kg Large – 22.5 kg Nothing to indicate infeasibility Return to Listing

Method of Matching Spirals has Errors Position <6000 km (1.5 % Earth-Moon Distance) Velocity ≈ 425 m/s Requires ≈ 13 kg Propellant Better trajectory matching requires more accurate model but Nothing to indicate infeasibility Trajectory Mismatch Return to Listing

Parking Orbit Selection Lower Orbit? Return to Listing

Parking Orbit Selection Higher Orbit? Return to Listing

Backup Slides Solomon Westerman Return to Listing

Total Cost ($M USD) Launch4.8 R&D2.3 Integration4.5 Purchase7.2 Overhead9.1 Total27.9 GLXP Prize ($M USD) Grand Prize20.0 Lunar Night5.0 Total25.0 Lose $2.9 M in 2012 USD! Return to Listing

Costing Model Differences 1.Overhead – 100g, 10kg k each 3 STK license, 15 MATLAB license – Arbitrary k each 10 STK license, 75 MATLAB license 2.R&D – 100g, 10kg k salary each + 50k per month equipment increases reliability by 2% per month – Arbitrary k salary each + 50k per month equipment increases reliability by 2% per month 3.Integration – 100g, 10kg $10k / kilogram – Arbitrary $10k / kilogram Return to Listing

Backup Slides Brad Appel Return to Listing

Electric Propulsion System Setup S From PCDU S/C Communication 9 9 Xenon System Thermal System Power / Intercomm 0.2 m No redundancies, no integration costs Return to Listing

Electric Propulsion System Specifications Specifications for the Hall Thruster – 100g Mission VariableValueUnits Thrust78.5mN Specific Impulse1950s Mass Flow Rate4.1mg/s Power Input1526W Efficiency Input Voltage350VDC Mass5.7kg Propulsion System Totals – 10kg Mission VariableValueUnits Wet Mass215kg Dry Mass30kg Required Power2,043Watts Burn time365days Thrust104mN Specific Impulse1964s Mass flow Rate5.4mg/s Specifications for the BHT-8000 Hall Thruster – Large Mission VariableValueUnits Thrust424mN Specific Impulse2250s Mass Flow Rate19.2mg/s Power Input7,600W Efficiency Mass25kg Propulsion System Totals - Large Mission VariableValueUnits Wet Mass3,810kg Dry Mass520kg Required Power38,773Watts Burn time365days Payload Capability4,545kg Return to Listing

LOx/LH2 would require an extra 600 kg, costing an extra $2.6M An ion thruster could accomplish the mission, but would require much more power than the HET Current technology places HET lifetime over 1 year Other Propulsion Options Return to Listing

Xenon Storage Thermal Analysis Allowed temperature path of propellant Maximize storage pressure for volume efficiency (~ 150 bar) Maintain tank temperature for gaseous Xenon phase:  Balance heat due to radiation and pressure drop with a 5 watt resistance heater Curve data from National Institute of Standards and Technology Temperature (K) Return to Listing