SCHEDULE. Presentation Schedule

Transcription

1 PROJECT ARTEMIS AAE 450 WEEK 05 PRESENTATIONS 2/13/2014 1

2 SCHEDULE 8:32 Parth S. 8:40 Krista G. 8:46 Michael C. 8:52 Jose Miguel B. - End Morning Session 10:32 Spenser G. 10:38 Ryan A. 10:44 Hani K. 10:50 Ben F. 10:56 Jessica C. Break 11:12 Eric M. 11:18 Cameron H. 11:24 Erik S. 11:30 Andrew E. 11:36 Arika A. Break 11:52 Finu L. 11:58 Bryan F. 12:04 Eric F. 12:10 Tas Powis 12:16 Divinaa B. 12:22 Joe A. 2 Presentation Schedule

3 PARTH SHAH APM INITIAL RISK ASSESSMENT 2/13/2014 o SINGLE LAUNCH LV RELIABILITY o GETTING INTO LEO o LUNAR ORBIT INSERTION 3

4 MISSION SURVIVABILITY FOR KEY EVENTS Events/Phases Probability of Success 1. Single Launch Capability of LV ~ Flight from Earth to LEO ~ 0.990(0.658) 3. LEO to Low Lunar Orbit (LLO) [2] ~ Landing on the Moon Living on the Moon Launch from Moon LLO to LEO [2] ~ Earth Re-entry + Landing --- PHASE 1 Mission Requirements: >90% colonists mission success >95% returning crew home safe Currently analyzing Phases 1-3, 7 Phase 3 = Phase 7 Most data from launch history reports, NASA logs Phase 4 and 6 has human element Analyzing Phase 5 in detail with Event Tree Defining success and catastrophic events Launch Vehicle Successes Attempts No. Launches Single Launch LV Single Launch Overall (klv) (nlv) Planned Reliability (plv) LV Failure Reliability Atlas V (401) Falcon SLS (Based on Saturn V) Falcon XX Bayesian 1 st order Probability of Success 4 Parth Shah APM

5 MISSION SURVIVABILITY FOR KEY EVENTS PHASE 2 Failure to Reach LEO P(fail) = Launch Failure Missed P(fail) = Launch Window P(fail) = Guidance and Navigation Issues P(fail) = 0.03 LV Single Launch Failure Rate P(fail) = 0.1 From LV Reliability P(fail) = Launch Process Issues P(fail) = Rocket Preparation Delaying Launch Weather Issues Delaying Launch P(fail) = 0.3 Control Systems Failure P(fail) = 0.01 Trajectory Correction Failure P(fail) = 0.01 Collision Avoidance Failure P(fail) = 0.01

6 KRISTA GARRETT MISSION DESIGN TRAJECTORIES 2/13/2014 o CREW MOON LANDER ASCENT PROFILE o TRANS-EARTH INJECTION ΔV o OTHER WORK: RADIATION EXPOSURE IN UNSHIELDED ROVERS 7

7 CREW MOON LANDER ASCENT PROFILE Vertical rise to 10 km Two-Point Boundary-Value Problem Minimal time = minimal propellant Reach desired altitude and velocity Results: Time for lunar ascent: 5 minutes, 10 seconds Propellant needed: 7.74 Mg Initial mass for ascent: Mg 8 Krista Garrett Mission Design

8 TRANS-EARTH INJECTION ΔV Δv = km/s Re-entry parameters: Altitude = km Velocity = km/s Flight Path Angle = Patched Conics Return (Krista Garrett) Hyperbolic Trajectory at Moon (Krista Garrett) 9 Krista Garrett Mission Design

9 MICHAEL CREECH MISSION DESIGN LUNAR OUTPOSTS 2/13/2014 o OUTPOST LOCATIONS o ENERGY COSTS FOR TRAVERSING THE LUNAR SURFACE 11

10 OUTPOST LOCATIONS 7 outposts per colony Both mare and craters Selected for optimal path between habitats and checkpoints 12 Michael Creech Mission Design

11 Power [kw] MISSION PROFILE AND ENERGY Mission profile for a rover path 3D view of mission profile Power profile and total energy requirement Example Shackleton to checkpoint 3 Total Energy = MJ Mission Profile - Power Map Horizontal Distance [km] 13 Michael Creech Mission Design

12 JOSE MIGUEL BLANCO MISSION DESIGN LOW THRUST TRAJECTORIES 2/13/2014 o LOW THRUST TRAJECTORY UPDATE, PATCHED TWO BODY MOTION IMPLEMENTED o MASS, POWER AND VOLUME FOR CARGO VEHICLE POWERED BY ELECTRIC PROPULSION 26

13 MASS POWER AND VOLUME OF CARGO VEHICLE wer simulation Mass vs mission time volume 27 Jose Miguel Blanco / Mission Design / Cargo and Lander

14 CARGO MASS TO LUNAR SURFACE mass volume Mass on lunar surface 28 Jose Miguel Blanco / Mission Design / Cargo and Lander

15 PROJECT ARTEMIS AAE 450 WEEK 05 PRESENTATIONS 2/13/

16 SCHEDULE 8:32 Parth S. 8:40 Krista G. 8:46 Michael C. 8:52 Jose Miguel B. - End Morning Session 10:32 Spenser G. 10:38 Ryan A. 10:44 Hani K. 10:50 Ben F. 10:56 Jessica C. Break 11:12 Eric M. 11:18 Cameron H. 11:24 Erik S. 11:30 Andrew E. 11:36 Arika A. Break 11:52 Finu L. 11:58 Bryan F. 12:04 Eric F. 12:10 Tas Powis 12:16 Divinaa B. 12:22 Joe A. 40 Presentation Schedule

17 SPENSER GUERIN CONTROLS CARGO VEHICLE 2/13/2014 o ELECTRONIC CONTROL SYSTEM SIZE o REACTION CONTROL THRUSTER SIZING 41

18 ELECTRONIC CONTROL SYSTEM Computing power still tentative. COMPONENT POWER [W] MASS [kg] VOLUME [] OPERATING TEMPERATURES [ C] Inertial Measurement Unit (IMU) Star Tracker (x2) , , Altimeter ,000 N/A Computing ,000 N/A Total , Spenser Guerin Controls

19 REACTION CONTROL THRUSTER SIZING Allocate 50 m/s for trajectory and attitude control and 20 m/s for lunar orbit insertion clean up. Isp = 220 sec m_pl = 40 Mg m_p = 1.3 Mg 43 Spenser Guerin Controls

20 RYAN ALLEN CONTROLS CREW TRANSPORT VEHICLE 2/13/2014 o HUMAN MOON LANDER SPECIFICATIONS o RADIO INTERFEROMETER ARRAY 44

21 HUMAN MOON LANDER SPECS Human Moon Lander Specification Mass to LEO (crew included) Power Required Value Mg W Volume to LEO 560 m 3 CATIA Model from Scott Sylvester Human Lunar Lander Navigation System: Navigation System Total Value 1) Computer System Mass 2) Inertial Measurement Power Unit (IMU) 3) 2x Star Tracker 4) Altimeter *Values from Spenser Guerin 45 Ryan Allen Controls *25.4 kg *150.8 W Volume *0.05 m 3

22 RADIO INTERFEROMETER ARRAY CUBE Field of 30 Radio Interferometer Array Cubes (RIA Cubes) Charge battery during lunar day using solar panels Operate during lunar night Communicate with L2 satellite Deployed by Heavy Rover 30 RIA Cubes Value Mass 3.0 Mg Power W* Figure based on Murchison Widefield Array Volume (packed) 30.0 m 3 *Value from Eric Menke 46 Ryan Allen Controls

23 HANI KIM HUMAN FACTOR HEAVY & LIGHT ROVER 2/13/2014 o LIGHT ROVER AND HABITAT ENERGY ESTIMATION

24 HEAVY & LIGHT ROVER Cabin Oxygen Tank Material Aluminum O m 3 Empty cabin mass Mg N m 3 Inside Volume m 3 Total Volume m 3 Material Volume m 3 Fig. 1:Light rover Dimension Fig. 2:Light rover body Fig. 3:Light rover inside 48 Hani Kim 2/12/2014

25 HUMAN FACTOR Burn up in the atmosphere (x) Microwave freeze drying 1~30 kw ~1 m 3 ~50 kg Light Rover (2 crews) (1) Hours of Available Weight(kg) Volume(m 3 ) Water Food total Weight of solid waste (4) Wet Weight Other waste : kg Dry weight kg/person/day Per colony for mission (kg) Heavy Rover (For 2 crews) (1) Hours of Available Weight(kg) Volume(m 3 ) Water Food total Heavy Rover (For 4 crews) (1) Hours of Available Weight(kg) Volume(m 3 ) Water 5, Food 5, total 5, Habitat Energy consumption (4) Energy provided : 88 kw Energy (kw) Lighting Water recycle (6) 2.5 Oxygen system (7) 1.5 Game room 3 Computer 0.5 Total Hani Kim 2/12/2014

26 BEN FISHMAN HUMAN FACTORS LIFE SUPPORT/OXYGEN 2/13/2014 o PRESSURIZED ROVERS 52

27 HF: PRESSURIZED ROVER (LIGHT) Heavy Rover (Updated) Number of People Hrs Spent Mass O2 [kg] Mass N2 [kg] Total Mass [kg] Volume O2 [m^3] Volume N2 [m^3] Total Volume [m^3] (emergency) Light Rover (First iteration) Number of People Hrs Spent Mass O2 [kg] Mass N2 [kg] Total Mass [kg] Volume O2 [m^3] Volume N2 [m^3] Total Volume [m^3] Power = 120 W/tank kept at 3000 psi 53 Ben Fishman Human Factors

28 HF: CTV OXYGEN/FIGURES Days Mass (O2) [kg] Mass (N2) [kg] Total Mass [kg] Volume (O2) [m^3] Volume (N2) [m^3] Total Volume [m^3] Power = 120 W/tank kept at 3000 psi Still working on placement/storage of oxygen on CTV Starting work on Habitat oxygen Photo credit (teammate): HaNi Kim 54 Ben Fishman Human Factors

29 JESSICA CALLINAN HUMAN FACTORS RESUPPLY MISSIONS 2/13/14 o RESUPPLY MISSION MASS AND VOLUME OPTIONS o CLOTHING VOLUME UPDATE o MEDICAL DEVICES o TYPES OF PEOPLE TO SEND 58

30 RESUPPLY MISSIONS Multiple options for resupply per colony Resupply 3 months 6 months* 9 months 12 months Food mass (Mg) Water mass (Mg)** Total mass (Mg) Food volume (m 3 ) Water volume (m 3 ) Total volume (m 3 ) * The optimal resupply is 6 months for food, as food losses its nutrients drastically after one year of storage. However, water can stay for the entire mission. [7] **Water mass after initial Mg and m 3 are there 59 Jessica Callinan Human Factors

31 CLOTHING MEDICAL PEOPLE Clothing per colony for 2 week cycles vacuum packed [1] Volume (m 3 ): Mass (kg): Medical devices IntraVenous Fluid Generation (IVGEN) [6][8][4] m 3 [2] Advanced Life Support Pack (ALSP) [2] m 3 Reference International Space Station Integrated Medical Group Medical Operations Book (ISS IMG) [3][5] People Assuming all have at minimum Masters degree in their field 2 engineers, 2 doctors, 2 scientists, 2 psychologists Figure: Advanced Life Support Pack, Jessica Callinan 60 Jessica Callinan Human Factors

32 ERIC MENKE COMMUNICATIONS VEHICLE SATELLITE DISHES 2/13/2014 o DISH SIZING o NEW HEAVY ROVER DIMENSIONS 69

33 VEHICLE DISHES Link Budget L2 Satellite Vehicle to L2 Interferometer to L2 Unit Carrier Frequency GHz Transmitter Power Watt Transmitter Line Loss db Transmitter Antenna Beamwidth deg Transmitter Antenna Pointing Offset deg Distance Between Transmitter and Receiver km Receiver Antenna Diameter m Receiver Antenna Beamwidth deg Receiver Antenna Pointing Error deg Desired Data Rate 3.00E E E+05 bps Bit Error Rate 1.00E E E-05 probability Implementation Loss db Transmitter Antenna Diameter m Gain Margin db Dish Volume E-05 m 3 Dish Mass kg 70 Eric Menke Comm

34 HEAVY ROVER Cabin Volume 3 m 3 Shielding Volume m 3 Total Volume m 3 Battery Mass Mg Empty Shielding Mass Mg Total Mass Mg 71 Eric Menke Comm

35 CAMERON HORTON AERODYNAMICS SAMPLE CARRIER RE-ENTRY 2/13/2014 o SCIENCE SAMPLE CARRIER LAUNCH PLAN o POD DESIGN AND SPECS o RE-ENTRY CALCULATIONS 73

36 SAMPLE CARRIER LAUNCH PLAN Mass of Sample (kg) Total Launches Launches per Year Launches per Colony (per Year) Total Volume (m 3 ) Pod Empty Mass (Mg) Pod Final Mass (Mg) Pod Volume (m 3 ) Frontal Surface Area (m 2 ) All renderings done in CATIA 74 Cameron Horton Aerodynamics

37 RE-ENTRY Heat Load, Heat Flux, and G vs. Entry Angle X: -8.1 Y: X: -6.5 Y: 1 Heat Load Heat Flux G Atmosphere Deflection 0.4 X: Y: entry angle gamma (degrees) Original code by Nicholas LaPiana Optimal Entry Angle Ballistic Coefficient Max Heat Rate (Watts/cm 2 ) Heat Load (Joule/cm 2 ) Max G Load (G) -8.1 o Cameron Horton Aerodynamics

38 ERIK SLETTEHAUGH STRUCTURES PAYLOAD CONFIGURATION 2/13/2014 o PAYLOAD CONFIGURATION o COSTS FOR LAUNCH VEHICLES 78

39 PAYLOAD CONFIGURATION Assume: Payload to Lunar Surface w/ SLS, Mass: 12Mg; Volume: 479.5m^3 Launch # Launch Identification #1 Communication Satellite #2 Communication Satellite #3 Communication Satellite Description Satellite at L1 #1 Satellite in Polar Obit, #2 Satellite in Polar Obit Satellite at L2 Launch Vehicle Delta IV (Small) Delta IV (Small) Delta IV (Small) All Bases Mass Volume Delta Mass Delta Volume Sleeping Area/Bathrooms 4 #1 Pod Habitat Eating/Meeting SLS Food Storage/Prep Lab/Medical 5 #2 Pod Habitat Rec Room SLS Initial Water, System & 6 #3 Habitat Cargo Storage ECLSS & Fire SLS Suppression Lunar Fission Reactor, Solar Panels, 7 #4 Habitat Cargo RTGs, 1 Way Point Battery, Laser SLS Communication Terminal Habitat Shield Support 8 #5 Habitat Cargo Structure, 3D Printing Machine, 1 Way Point SLS Battery 9 #6 Habitat Cargo 3 Way Point Batteries SLS #1 Heavy Rover Heavy Rover - Tires, Cab, etc. SLS #2 Heavy Rover Battery - Heavy Rover SLS #3 Heavy Rover Heavy Rover - Tires, Cab, etc. SLS #4 Heavy Rover Battery - Heavy Rover SLS #1 Light Rover Light Rover SLS #2 Light Rover Light Rover SLS #3 Light Rover Light Rover SLS #4 Light Rover Light Rover SLS Launch # Launch Identification 18 #1 Construction Vehicles 19 #2 Construction Vehicles 20 #3 Construction Vehicles 21 #1 Science Sample Carrier 22 #1 Crew Transport 23 #1 Resupply 21* 11* & 13* #1 Interferometer (Packed with #1 Science Sample Carrier) # 1 Skylight Repel System (1 Plate Packed with #2 & 1 Plate w/ #2 Heavy Rover) Description Soil Transport Vehicle, 1 Way Point Battery Launch Vehicle Mass Total # of Launches Delta Mass Delta Volume , Erik Slettehaugh Launch Vehicle & Com Sats Volume Delta Mass Delta Volume SLS Bulldozer SLS Helper Robots SLS Sample Capsule, Rail Gun, etc., 1 Way Point Battery SLS Crew Transport Vehicle, Crew Moon SLS Lander Food, Water, Personal Items, Spare Parts SLS Shackleton Base 30 Radio Interferometer Array Cubes Skylight Repel System, 2 Elevator Plates Skylight Base SLS N/A N/A SLS N/A N/A

40 COST PER LAUNCH Launch Vehicle Type Payload Fairing Payload Volume (m^3) Assisted by: Finu Lukose, Sadie Holbert Mass (LEO) (Mg) Cost ($1M)/Mg Cost ($1M)/(m^3) Cost / Launch ($1M) SLS 8.4m Diameter $ 3.85 $ 0.47 $ m Diameter $ 3.85 $ 0.33 $ Falcon Heavy Composite Fairing $ 2.74 $ 0.52 $ Dragon Spacecraft Falcon 9 & Trunk $ 4.30 $ 0.73 $56.50 Composite Fairing $ 9.51 $ 0.45 $ Large Payload Fairing (LPF) $ 9.18 $ 0.32 $90.00 Extended Payload Fairing (EPF) $ 7.90 $ 0.30 $95.00 Atlas V Extra Extended Payload Fairing $ 6.55 $ 0.28 $ (XEPF) 431 Short $ 8.67 $ 0.18 $85.00 Medium $ 7.04 $ 0.14 $95.00 Long $ 6.90 $ 0.14 $ Small Payload Fairing $ 9.79 $ 0.26 $90.00 Delta IV Medium Payload Fairing $ 9.04 $ 0.17 $ Large Payload Fairing $ 4.95 $ 0.16 $ Assisted by: Arika Armstrong 80 Erik Slettehaugh Launch Vehicle & Com Sats SLS Payload Fairing Unused Space Cargo Transport Vehicle Payload Bay 10m 9m 9.738m 3.6m 9.8m 6.438m 19.1m

41 ANDREW EMANS STRUCTURES HABITAT SHIELDING 02/13/2014 o HABITAT SHIELDING SUPPORT STRUCTURE o 3D PRINTING WITH REGOLITH 84

42 HABITAT SHIELDING Initial analysis shows a reduced launch mass with the use of a 3D lunar concrete printer instead of carbon fiber as the shield support material. Pros/cons for concrete Requires a 3D printer larger than any ever built which could fail and leave the habitat unshielded Reduces initial mass from Earth by ~30% compared to carbon fiber (saves ~84% mass for additional check-points) Printer can be reused for check-point construction Pros/cons for carbon fiber Easier to install Reliable since more than one robot can complete the task of installing 85 Andrew Emans Structures

43 MATERIAL NUMBERS Lunarcrete Total Mass [Mg] Volume [m^3] Power [kw] 60 0 Requires: 3D Printer Gantry Helper Robot Carbon Fiber Construction Robot Note: Mass and volume values were calculated by assuming thin walls and the primary mode of failure is due to buckling. Sketches drawn by Andrew Emans, except for the three arcs in the top sketch which were skillfully drawn by Arika Armstrong 86 Andrew Emans Structures

44 ARIKA ARMSTRONG STRUCTURES CARGO SIZING 02/13/14 o CARGO POD o ALLOWABLE PAYLOAD MASS 90

45 CARGO SIZING Cargo Pod External Diameter [m] External Height [m] Internal Diameter [m] Internal Height [m] Mass [Mg] Lander Vehicle External Dimensions* [m] 9.8 D x 9 L Inert Mass [Mg] Max Landed Mass [Mg] Max Cargo with Pod [Mg] Propellant Mass [Mg] Arika Armstrong Structures

46 ADDITIONAL CARGO SYSTEMS Landing Strut System Based on Apollo 3% ~ Mg for landing gear Marman Clamp ~0.1 Mg for a clamp 92 Arika Armstrong Structures

47 FINU LUKOSE PROPULSION THRUSTER SIZING & CHASSIS DESIGN 2/13/2014 o THRUSTER SIZING AND PROPELLANT MASS o CHOICES FOR THRUSTERS o PROGRESS ON CHASSIS DESIGN 93

48 COM SAT PROPULSION SIZING Basic Forces Model on cylindrical satellite Refined model for greater maneuverability/pointing capability Code estimating thrust necessary for maneuvers (Controls Team) Sizing of thrusters based on thrust Off the shelf solutions Modified AAE 539 code 94 Finu Lukose Propulsion

49 CHOICES FOR THRUSTERS Controls sizing for micro-thrusters to resist various environmental torques (~5E-5N) Approximate delta V for station keeping ~750 m/s for 4.5 year (based on LRO) Bi-Propellant Thrusters Model Thrust (N) Propellant Isp (s) System Mass (kg) Propellant Mass (kg) Total mass (kg) Total Volume (m^3) R-6D 22 MMH/NTO R-1E 110 MMH/NTO R-4D 490 MMH/NTO Model Thrust (N) Propellant Mono-Propellant Thrusters Isp (s) System Mass (kg) Propellant Mass Total mass (kg) Total Volume (m^3) MR-103D 1 Hydrazine MR-111C 4 Hydrazine MR-106E 22 Hydrazine Finu Lukose Propulsion

50 BRYAN FOSTER PROPULSION SCIENCE SAMPLE RETURN 2/13/2014 o ELECTROMAGNETIC LAUNCHERS o BIPROPELLANT ROCKET 99

51 RAILGUN CONCEPT Sample Mass (Mg) Sample and Capsule Mass (Mg) Force (kn) Current (ka) Voltage (V) Energy (MJ) Railgun Mass (Mg) Railgun Volume (m^3) Launches Ballistic Return Velocity m/s 100 Bryan Foster Science Sample Return

52 SAMPLE RETURN ROCKET Sample Mass (Mg) Sample and Capsule Mass (Mg) Dry Mass of Rocket (Mg) Propellant Mass (Mg) Gross Mass of Rocket (Mg) Volume of Rocket (m^3) Airbus Aestus Rocket Engine MMH/N2O4 propellant Uses same low energy ballistic trajectory Releases samples and burns up in reentry. 101 Bryan Foster Science Sample Return

53 ERIC A FLORES / PROPULSION CREW TRANSPORT VEHICLE 2/13/14 o CLM PROPELLANT ANALYSIS UPDATE o CTV PROPELLANT ANALYSIS UPDATE

54 CLM PROPELLANT ANALYSIS UPDATE Hovering Capabilities (Descent Stage) 60 Seconds = 0.27 Mg of Propellant o Assuming complete use of 25 Mg before hovering. ***Hovering Code Credit to Sean Snoke Vehicle Propellant Mass [Mg] Propellant Volume [m^3] LM Ascent LM Descent ***Propellant Mass Credit to Krista Garret 112

55 Volume (m^3) CTV PROPELLANT ANALYSIS UPDATE Trajectory Propellant Mass [Mg] Propellant Volume [m^3] LEO to LLO LLO to LEO Total CTV Propellant for the Mission J-2 (Old Estimates) J-2X (New Estimates) Total Selected Engines Optimization Code Bringing Costs vs. Sending Costs? 50.7% Less Propellant 113

56 ANDREW POWIS POWER/THERMAL ROVER POWER SYSTEMS 2/13/2014 o VEHICLE TEAM LEADER: PRESSURIZED ROVERS o FUEL CELLS TO BATTERIES o ROVER SPECIFICATIONS 117

57 TRANSITION FROM FUEL CELLS TO BATTERIES Battery as shielding on Heavy Rover. Decrease in final mass. Minimal change to launch mass. Avoids the hydrogen storage issue. Power system commonality between rovers and construction bots. Easier to achieve maximum power output. [3] Battery/ Shielding Cabin 118 Eric Menke Andrew Powis Power/Thermal

58 Return Journey LIGHT ROVER / ROVER SUMMARY 1.8 m Ha Ni Kim Life Support Systems Window Light Rover Cabin 1.8 m Standing Room 0 km 100 km Habitat 3.5 m Rover Heavy Light Maximum Power (kw) Total Energy (kwh) Battery Mass (Mg) [3] Battery Volume (m 3 ) [3] Number of Battery Cells [3] Rover Launch Mass (Mg) Rover Launch Volume (m 3 ) Rover Final Mass (Mg) Rover Final Volume (m 3 ) Checkpoints 750 km Journey Out 119 Andrew Powis Power/Thermal

59 DIVINAA BURDER POWER&THERMAL CTV POWER SYSTEMS 2/13/2014 o SOLAR ARRAYS o LITHIUM-ION BATTERIES 120

60 SOLAR ARRAYS Stored in Service Module Ultra-flex Butterfly wing shape (based on Orion) Power: 15 kw Volume: 30 m^3 Advantages: Low Mass Compact storage Durable 5.5 m 121 Divinaa Burder Power & Thermal

61 LITHIUM-ION BATTERIES & SYSTEM DIAGRAM Lithium Ion Batteries Re-chargeable Secondary power Dark Periods Launch operations Re-entry Power: 35 kwh Mass: 400 kg (with insulation) Volume: 0.18m^3 Battery Battery Secondary System Power Bus PCU Solar Arrays Power Bus PCU Primary System Battery Battery Power Systems Structure Solar Arrays Lithium-Ion Batteries 2 Power Buses 2 Power Control Units (PCUs) Sub Systems Sub Systems 122 Divinaa Burder Power & Thermal

62 JOSEPH AVELLANO PT/CARGO CARGO VEHICLE THERMAL 2/13/2014 o OPERATING THERMAL TEMPERATURES o STORAGE OF ELECTRONIC SYSTEMS o MULTI LAYER INSULATION 126

63 OPERATIONAL TEMPERATURES Low Earth Orbit Temps: -75 C 65 C Operational Temp for Battery: Lithium Ion (NCA) - 0 C Total volume of systems :.081 m 3 52 cm x 52 cm x 38 cm carbon fiber container 1 cm thick 127

64 SIZING Carbon Fiber Container: 52 cm x 52 cm x 38 cm 1 cm thick Multi Layer Insulation: cm x cm x cm 5.08 cm thick Carbon Fiber Container Multi Layer Insulation Mass (kg) Power (W) Volume () Heating Pad E-6 Thermal Control Unit E-6 128