1. Justin Treptow Alex Morrese Alexis Mendez Casselle Russell John Klaus Robert Cooper Thilina Fernando Zoe Morozko Faculty Advisors: Dr. Dan Kirk Greg Peebles Paul Martin Ben Burnett Shriman Shiva Damian Harasiuk 1

2.  Develop “Mission Plan” with in 4 years, minimum expense, with no military or heritage hardware  Establish a program that continues beyond our graduation  Bring useful talent into the program such as: Universities, Professionals, Graduates 2

3.  Land in GLXP approved landing site near a Historical Artifact  Traverse 500m in a deliberate manner around a Historic Artifact  Transmit minimum dataset to earth  On arrival Mooncast > 500MB  Mission Complete Mooncast > 500MB 3

4. On Arrival Mooncast Descent Video > 2min. Near Real Time (NRT) transmitted ASAP Camera shot vertically or horizontally Arrival Video (Landing Site and Environment) > 30sec NRT transmitted ASAP Post Arrival Images: Panorama, 3 self portraits depicting 40% of craft, GLXP Logo Cluster and Lunar surface Transmit video message, email, text message NRT video > 6 min, shot from various angles HD video > 6 min, shot from various angles 4

5. Mission Complete Mooncast  NRT departure video > 30 sec, transmitted ASAP, shot of transition onto lunar surface  NRT looking back video > 30 sec, transmitted ASAP, shot of original landing site as vehicle moves away  Looking back photo depicting landing location  NRT Mid-Journey video > 30 sec, transmitted ASAP, shot while transversing  Mid-Journey photo at 250 m from landing site depicting > 40% of rover surface  Journey’s End Panoramic at 500m with lander/landing sight in view  Journey’s End Self Portrait portraying > 40% of transversing vehicle  3 Journey’s End photos of logo cluster and moonscape  NRT video > 5min, shot from various angles  HD video > 5 min, shot from various angles 5

6. Design Requirements System System Drivers (Metrics) Analysis Trade Study Design Conflicts Big Picture System Level Subsystem Level Interface Level System Drivers (Metrics) Analysis Trade StudySub -System System Drivers (Metrics) Analysis Trade StudyInterface 6

7. RequirementsPersonnel Allocation Landing in GLXP approved landing siteZoe, Rob, John, Alex, Justin, Damian, Paul, Ben Traverse 500mThilina, Alexis, Shriman Transmit Minimum DatasetCasselle Note: There are no walls in our group Project Start: 5 Seniors Junior Design: 8 Seniors Senior Design: 11 Members ( 8 seniors & 3 graduate level members) Senior Design Current: 12 members (8 seniors & 4 graduate level members) 7

8. Launch Vehicle Trade Study 8 PropertiesDelta IIAtlas V - 401Zenit 3SLFalcon 1Pegasus XL Dnepr Payload to LEO (kg)~6,0969,750~7,0007234543700 Cost (millions of dollars)5075707169.5 Reliability (success rate, %)99%100%97%0%87%97% Launch Location (distance to equator, km) ~3165 km (Cape) Equator (0 km) ~3165 km (Cape) The sky Baikonur, Kazakhstan Weight %Scores Payload to LEO (kg)0.44108316 Cost (millions of dollars)0.355121089 Reliability (success rate, %)0.29108178 Launch Location (dist. to equator, km)0.0588108 7 1.643.21.20.42.4 1.750.350.73.52.83.15 1.821.60.21.41.6 0.4 0.50.40.50.35 Overall5.556.7565.35.1 7.5

9.  Converted Soviet ICBM  97 % success rate as ICBM  3 Stages  Payload to LEO: 3700 kg  Cost: $9.5 million  Launches out of Kazakhstan at different inclinations http://snebulos.mit.edu/projects/reference/launch_vehicles/DNEPR/Dnepr_User_Guide.pdf 9

10. Inclination = 50.5 Inclination =64.5 Inclination = 87.3 Inclination = 98 10 Inclination (i) Orbit inclination with respect to earth’s equator

11.  92% Drop in Payload Mass over Orbit Altitude Increase of 600 km (i = 50.5°) Dnepr Users Guide pg 19 11

12.  Hohmann Transfer is the preferred method  Ion propulsion’s time scale is outside of mission requirements  SMART -1 ion propelled mission took 13 months 12

13.  Purpose To develop a analytic tool that accelerated analysis or Hohmann Transfers  Input (Design Specifications) Orbit altitude ISP Mass to be at final location  Output (System Drivers) Total & Incremental ΔV Total Propellant Mass Required Mass Ratio 13 Δv: velocity change required to alter the orbit of a space craft

14.  From these results: ▪ Velocity can be determined at a specific point along the trajectory ▪ Mass of the fuel required will aid in tank design requirements ▪ Mass available for landing will allow for payload mass estimates 14

15.  Acquired Satellite Tool Kit (STK)  Using the Astrogator Suite  $100,000.00 value  Utilized by: Lockheed Martin, Northrop Grumman, NORAD, Florida Institute of Technology (Dana Carmody), ect.  Paul Martin is taking point with the simulation with oversight from senior design team members Capacity to model complete mission simulation Launch – Mission Complete Images provided by AGI.com 15

16.  What orbit altitude corresponds to most efficient use of Dnepr’s lifting capabilities and propellant in spacecraft?  ΔV variations occur when transferring from a circular earth orbit to an elliptical earth orbit with apoapsis at L1 point. 300km 600km 900km L1: location where gravitational force from earth and moon are equal L1 16

17. Orbit Altitude (km) Circular Earth Orbit Velocity (km/s) Required Velocity at Periapsis of Elliptic Orbit (km/s) ΔV (km/s) Dnepr Mass Delivered (kg) 3007.7310.813.583700 4007.6710.733.063300 5007.6110.653.042800 6007.5610.573.011900 7007.5010.503.001200 8007.4510.422.97700 9007.4010.352.95300 Maximum ΔV difference of 0.63km/s, 17.6% change 17

18. Large change in orbiting mass for small change in ΔV. Best to use Dnepr to launch between 3300kg and 3700kg mass into orbit 18

19. Fuel Specific Volume (m3/kg)Oxidizer Specific Volume (m3/kg) Mixture Ratio (Ox:Fuel)Isp (s)MR Liquid Bi- Propellants Hydrogen1.41E-02Oxygen8.77E-0453904.12 Kerosene1.34E-03Oxygen8.77E-042.293016.25 Hydrazine9.96E-04Oxygen8.77E-040.743135.83 MMH1.15E-03Nitrogen Tetroxide6.90E-041.732807.18 UDMH1.26E-03Nitrogen Tetroxide6.90E-042.12777.33 Liquid Mono-propellants Hydrazine9.96E-04- --19916.00 Hydrogen Peroxide6.94E-04- --16528.34 Solid Propellants Aluminum+HTPB Ammonium Perchlorate5.12E-042.122677.90 Aluminum+PBAN Ammonium Perchlorate5.12E-042.332667.96 Mass Ratio based on total Δv of 5.42km/s Spacecraft System Engineering. 3rd ed. Fortescue, P. pg 183 http://www.braeunig.us/space/propel.htm http://www.astronautix.com/ ROCKET PROPELLANTS, Warren FA, Reinhold Pubising 1958 http://www.braeunig.us/space/propel.htm http://www.astronautix.com/ 19

20. Initial Tank Volume Required (m 3 ) % Volume Required FuelOxidizer Mixture Ratio (Oxidizer:Fuel) Specific Impulse (s)MRFUELOXIDIZERTOTAL Standard Payload Module HydrogenOxygen53904.126.752.108.8667.6 KeroseneOxygen2.293016.251.256.187.4356.57 HydrazineOxygen0.743135.831.244.085.3240.6 MMH Nitrogen Tetroxide1.732807.181.353.815.1739.4 UDMH Nitrogen Tetroxide2.12777.331.324.686.0045.8 Standard Payload Module Volume: ~13m 3 20

21. For equipment operating on Earth's surface under ideal operating conditions. For equipment operating on Earth's surface under ideal operating conditions. Specific cells cited are not space rated. Specific cells cited are not space rated. Only hydrogen reactant weight included, oxygen requirement has been neglected. Only hydrogen reactant weight included, oxygen requirement has been neglected. All cells use manufacturer's stated output. All cells use manufacturer's stated output. SourcePower OutputMassVolumeReactant usageOp. Temp.Reactant Pressure PEM fuel cell1.2-2 kW44-110kg0.128 – 0.192 m32.14 – 4.69 m3/hr2 – 46 C125 - 142 Kpa alkaline fuel cell6kW200kg0.553 m34.3 m3/hr-20 - 40 C400-600 Kpa solar panels205W17.96kg.0734 m3n/a Li-ion batteries300-1500W/kg576 J/g972 J/m3n/a RTG330W55.5 kgnone stated4.5kg4400Wn/a 21 Sourcereactant/unit#unitstotal mass req.total volume req. PEM fuel cell25.7987525678.993760.96 alkaline fuel cell90.939841290.939840.553 solar panelsn/a29.26829525.65853662.148292683 li-ion batteries36006.66666724006.666672133333.333

22. 1.Alkaline Fuel Cell  Used on Apollo and STS missions  High performance compared  Rarely sees modern design  Prone to carbon dioxide saturation 2.Polymer Electrolyte Membrane Fuel Cell  Current industry standard  Has never been space rated  Lower performance than alkaline cells  Highly sensitive to fuel impurities 3.Photo Voltaic Cells (Solar Panels)  Widely used on previous missions  Power generation is time independent  Requires deployment system  Large volume requirement 22

23. Site Latitude Longitude Comments Apollo 11 Landing1N24EFirst manned moon mission Apollo 12 Landing3S24W Apollo 12 LEM impact4S21W Apollo 13 Saturn IVB impact3S19WOnly mission success during Apollo 13 Apollo 14 Landing4S18WAllan Shepherd plays golf Apollo 14 LEM impact3S20W Apollo 14 S-IVB impact8S26W Apollo 15 Landing26N5EUsed modular equipment transporter Apollo 15 LEM impact26N1E Apollo 15 S-IVB impact2S11W Apollo 16 Landing9S16EFirst mission to use the lunar roving vehicle Apollo 16 S-IVB impact1N24W Apollo 17 Landing21N31ELast manned moon mission Apollo 17 LEM impact21N31EUsed lunar roving vehicle Apollo 17 S-IVB impact5S13W 23

24.  The goal has been made; land at the location of the first Apollo manned Moon landing, Apollo 11 Luna 20 Luna 16 24

25. Qualifies as a historical landing spot which accomplishes one of the extra goals presented by the X-Prize competition Location of “Sea of Tranquility”: ▪ Proven successful landing location ▪ relatively rock and obstruction free Equatorial location requires less fuel for orbital corrections and an easier trajectory Equatorial landing trajectory between ±10º Latitude provides the most possible landing sites in case we “miss” 25

26. MotorsManufactureModelPowerS/R Drive motorsMaxon MotorsRE-max 2915W 48V 0.283A 1.59A (Starting ) Upon Request Steering motorsMaxon MotorsRE-max 216W 48V 0.151A 0.598A(starting) Upon Request Cameras HD11.16 W (max estimation) No SD=16.8 W (max estimation) No Tentative power budget based on selection of components. Research into cameras and other onboard electronics is on going. Given this, a total consumption of 150 watts will be estimated. This is based on the power consumption of previous rovers that have been designed. 26

27. Comparable Systems: Spirit & Opportunity both operated at 150W SpectroLab’s Triple Junction Solar Cell 316 W/m 2 Design Recommendation: Power consumption <=150W 27 NASA Ideal Power Generated by Solar Panels Triple Junction Solar Panels = 0.474 m 2 surface area

28. Lunar Conditions ≠ Ideal Direct solar Radiation Panels above normal operating conditions, Loose efficiency. Assumptions:  Solar Radiation Sc = 1353 W/m 2  Emissivity of Solar Panels0.85  Temp Lunar Environment 107 ⁰C Analytical Analysis (Radiation Transfer)  Panel Surface Temp Ts = 197°C  Voltage drop 1.13v/cm 2  Negligible Current Increase 1.24*10 -4 Amp These losses will have to be compensated by including extra Solar Cells 28

29. Parameters that determine performance:  Radiation Susceptibility  Power Consumption  Operation Temperatures  Mass  Resolution  Frame Rate 29

30. 30 Requirements for HD camera  720p resolution 1280x720 pixels  15 frames a sec minimum Requirements for SD camera  Near real time transmission  15 frames a sec minimum  Always on

31. CAMERACOST ($)POWEROPERATING TEMPMASS BODYADAPTBATT Canon EOS Digital Rebel XS Standard Definition CMOS Sensor 800 TYPE: Rechargeable Lithium Ion VOLTAGE: 7.4V DC CAPACITY: 1080mAh LIFE: 500-600 shots{NO FLASH] 400-500[50% FLASH] 0-40˚C (32-104 ˚F) 475g (16.8oz)80g (2.8oz) Canon VIXIA HF10 High Definition 1080i 810 SUPPLY: 100-240V AC,50/60Hz OUTPUT: 8.4V DC,1.5A 0-40˚C (32-104 ˚F) 380g (13.4oz)135g (4.8oz) Nikon D40 Standard Definition CCD 480 TYPE: Lithium Ion VOLTAGE:7.4V DC CAPACITY:1000mAh 0-40˚C (32-104 ˚F) 475g (17oz)51g (1.8oz) Panasonic HDC SD9 High Definition 1080i CCD 650 SUPPLY:110-240V AC, 50/60Hz OUTPUT: 9.3V DC, 1.2A 0-40˚C (32-104 ˚F) 275g (9.7oz)115g (4.0oz) Sony Alpha DSLR-A200 Standard Definition CCD 500 TYPE: Lithium Ion VOLTAGE: 8.4V DC, 2.0A CAPACITY:1650mAh 0-40˚C (32-104 ˚F) 532g (18.8oz)78g (2.8oz) Sony Alpha DSLR-A350 Standard Definition CCD 800 TYPE: Lithium Ion VOLTAGE: 8.4V DC, 2.0A CAPACITY: 1650mAh 0-40˚C (32-104 ˚F) 582g (20.5oz)78g (2.8oz) Sony HDR-HC9 High Definition 1080i CMOS 980 SUPPLY:100-240V AC, 50/60Hz OUTPUT: 8.4V DC, 0.35-0.18A 0-40˚C (32-104 ˚F) 550g (1lb 3oz)170g (6.0oz) Sony HDR-SR12 High Definition 1080i CMOS 1150SUPPLY: 100-240V AV, 50/60Hz OUTPUT:8.4V DC, 0.35-0.18A 0-40˚C (32-104 ˚F) 570g (1lb 4oz)170g (6.0oz) Camera Analysis 31

32.  Selene(Kaguya)  took HD images of lunar surface and earth setting on its horizon  16.5kg,50W, CCD sensor  ISS (Imaging Science Subsystem)  Cassini Orbiter- took images of Saturn  57.83kg,55.9W, CCD sensor  ITOS (NOAA)  Satellites used for meteorology  AVCS (Advanced Vidicon Camera System)- used before CCD 32

33. 1. Total Ionizing Dose (TID) Changes in threshold voltage 2. Displacement Damage (DD) Movement in Si lattice Effect functions (ex. Power) 3. Single Event Effects (SEE) Single Event Upset Single Event Latch up Single Event induced Burnout 33

34. CCD camera w/o RH CCD camera w/ RH t=0 (100Gy/hr) t=1hr (100Gy/hr) 34

35.  Radiation Hardening Techniques  Physical- ▪ Microchips that are made from insulating substances (Silicone Oxide & Sapphire) ▪ Shielding the package or the chips themselves (depleted Boron) ▪ Using components with a wide band gap  Logical- ▪ Error correction (parity check) ▪ Redundant elements 35

36. Option 1  Design Electronics with hardening in mind Option 2  Shield commercial electronics Case study will be performed to select 36

37. System (with 4 wheels) total # of Joints # springs and dampers Wheel travel Rocker Bogie2 differential joints 0Depends on design Independent suspension 24 (6 per wheel) 4 springs 4 dampers Limited by the spring or damper Source: Heiken et al, Lunar Source Book Cambridge university press 1991 System Drivers: Complexity, Mass, Vehicle Velocity Current Conclusion Rocker Bogie System 4 wheel Rocker Bogie system selected No springs or dampers Has the greatest wheel travel Least number of joints Proven history 37

38. Wheel Design Literature review of mechanical properties of lunar regolith Recommended ground pressure of 7-10 Kpa for wheeled vehicles 38

39.  Complete trafficability calculations  Wheel sinkage  Gross pull per Wheel  Soil resistance per wheel  Perform kinematic analysis on the Rocker Bogie and fine tune the design to fit our particular needs. 39

40. 40 Launch Vehicle Orbital DynamicsTransfer Vehicle Lander End Fall 08 Big Picture System Sub System Interface End Spring 09 2 Iterations 3 Iteration >2 Iterations 0 Iterations

41. 41 Big Picture System Sub System Interface Rover PowerImagingCommunications 2 Iterations 2 Iteration 1 Iterations End Fall 08 End Spring 09 Roving Vehicle 1 Iterations

42.  Document will contain  Requirements  System level Analysis ▪ System Drivers (performance determining equations) ▪ Evaluation of System Drivers ▪ Sub-System level (limited)  Document will contain  Requirements  System level Analysis  Sub-System level Analysis  Interface Analysis 42

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