1. Shackleton Crater Reconnaissance Mission PDR Trevor Fedie Jason Breeggemann Brian Evans Mike Gavanda Matt Gildner Jeromie Hamann Brian Nackerud Andrew Smude Jordan Stewart

2. Project Overview Image 5 km annular region around Shackleton Crater (at Moon’s south pole) Image 5 km annular region around Shackleton Crater (at Moon’s south pole) –3 color imagery –10 cm resolution –Image 80% of region in one month Act as communication relay for lunar lander exploring interior of crater Act as communication relay for lunar lander exploring interior of crater –2.5 Mbps S-band from lander –Transmit to Earth once per day

3. Project Overview Baseline orbit of 30 x 216 km Baseline orbit of 30 x 216 km –Period approximately 120 minutes –12 orbits each day Spacecraft must fit into Taurus Launch Vehicle Spacecraft must fit into Taurus Launch Vehicle –Goal of 445 kg for everyday launch opportunities –Must handle spin stabilized upper stage (60 rpm) –Must interface with launch vehicle upper stage –Must fit into launch vehicle envelope

4. Mission Profile Launch using Taurus Launch Vehicle Launch using Taurus Launch Vehicle De-spin, make burn for Earth-Moon transit De-spin, make burn for Earth-Moon transit Burn to enter Lunar orbit Burn to enter Lunar orbit Begin imaging region around crater Begin imaging region around crater Send data back to Earth once per day Send data back to Earth once per day Finish imaging crater in about 14 days Finish imaging crater in about 14 days Lunar lander operations begin 4 months after launch and last for one year Lunar lander operations begin 4 months after launch and last for one year

5. Spacecraft Data Flow Flight Computer SSR Transponders Star Tracker Camera Gyroscope Reaction Wheels Rocket Motor Thrusters Temp Sensors Antennas

6. Camera Crater Imaging Strategy (Note: not to scale)

7. Camera HiRISE (used on Mars Reconnaissance Orbiter) HiRISE (used on Mars Reconnaissance Orbiter) –Pushbroom TDI imager (4,048 pixels across swath) –0.5 m aperture –28 Gbits internal data storage –Internal LUT image compression –Mass: 65 kg –Average power: 60 W

8. Computing & Data Storage –3U Compact PCI  PowerPC RAD750  Enhanced Power PCI Bridge –SSR  P9 family  160 Gbits BOL  Built-in FELICS  256 Mb SDRAM

9. Communications Earth Comm. Earth Comm. –Cassegrain Antenna  D=.98m  d=0.3m  Power usage 40 W  Double reflection

10. Earth Communications Transmit high resolution photos Transmit high resolution photos –Only one contact with White Sands per day –Around 90 Gbits a day in pictures –Ka-band 26 Ghz parabolic dish –Data rate:100 Mbps –Power: 40 W –Gain: 44 db –White Sands receiving: 45 db/T Similar to LRO Similar to LRO

11. Rover Communications Relay for rover in crater Relay for rover in crater –Required data rate of 2.5 Mbs –Required S-band from rover –2.3 Ghz Omni-directional transmitter –Power: 5 W –Gain: 5 db –Rover Gain: 5 db  L3 T&C transceiver MSX-765 Store Data onboard satellite until downlink to White Sands ground Station Store Data onboard satellite until downlink to White Sands ground Station

12. Communications Earth Earth –4230 sec average daily window –Tracking  Azimuth: 0.01617 deg/s  Elevation: 0.0154 deg/s Moon Moon –12 communication windows  6.533 min over crater  117.6 Gbits of data a day  2 min required to send to earth

13. Pointing stability/accuracy Pointing stability/accuracy Torque disturbances must result in 10 cm or less ground displacements during exposure time Torque disturbances must result in 10 cm or less ground displacements during exposure time Attitude accuracy Attitude accuracy -Roll axis <2.5 arcmin -Pitch axis <8.6 arcmin -Yaw axis <2.7 arcmin Maneuverability Maneuverability 3-axis control 3-axis control Pointing reassignment as fast as 90 deg in 6 minutes Pointing reassignment as fast as 90 deg in 6 minutes Attitude Determination and Control

14. SED 16 Autonomous Star Tracker by Sodern SED 16 Autonomous Star Tracker by Sodern –36 arcsec in pitch/yaw, 108 arcsec roll (bias plus noise) – 10 Hz update –25x25 deg field of view –Mass: 2.9 kg (with Baffle) –Average power: 10.7 W

15. Attitude Determination and Control Scalable SIRU (gyro) by Northrop Grumman Scalable SIRU (gyro) by Northrop Grumman –Achieves Gyro Bias stability of 0.0003 deg/hr Four HRGs (Hemispherical Resonator Gyro), with associated loop control/readout/thermal control electronics, and sensing along the octahedral-tetrad axes –Low noise –Mass: 7.1 kg –Average power: 38 W

16. Attitude Determination and Control Momentum build up Momentum build up Disturbance Torques Disturbance Torques - Gravity gradients - Solar pressure - Internal - Deployables Slewing Maneuvers Slewing Maneuvers - Minimum twice a day HR14 Constellation Series Reaction Wheels by Honeywell HR14 Constellation Series Reaction Wheels by Honeywell –Max Reaction Torque 0.2 N-m –Momentum Capacity 50 N-m-s –Mass: 8.5 kg

17. Propulsion System Rocket Motor: -EADS Astrium S400-12 -MMH/MON-1 -420N Isp:318s -Total  V needed: 1011m/s

18. Propulsion System Twelve 4N thrusters Twelve 4N thrusters –MMH/MON-1 Fuel/Oxidizer –EADS Astrium S4 Two 0.05 cubic meter tanks Two 0.05 cubic meter tanks –Composite structure –3600 psi rated –Lincoln Composites

19. Determination of Thermal Environment Moon surface temp Moon surface temp Altitude and attitude Altitude and attitude Lunar view factor Lunar view factor Intense reflected IR from lunar surface. Intense reflected IR from lunar surface. Thus objects placement will important Thus objects placement will important If thermo enviroment is compromised anywhere and active system will be used If thermo enviroment is compromised anywhere and active system will be used

20. Thermal Analysis Equilibrium temperature range of 270-310 Kelvin Equilibrium temperature range of 270-310 Kelvin Electric heaters and sensors on items that do not fall within their ideal ranges. Electric heaters and sensors on items that do not fall within their ideal ranges. Use of heat tubes and radiators if needed Use of heat tubes and radiators if needed Will be tuned in by a dynamic model Will be tuned in by a dynamic model

21. Structural System Determination Properties of Ti an Al Properties of Ti an Al Bulk system of TI. Bulk system of TI. AL only used where conduction requires it. AL only used where conduction requires it. Current structural mass of approximately 18Kg. Current structural mass of approximately 18Kg. Titanium will also have less thermal expansion Titanium will also have less thermal expansion ALTi Yield Strength (MPa) 475- 520 480- 1170 Coefficient of Thermal expansion ≈238-11 Conductivity (W/m - K) 88- 210 6-17 Density(Mg/m³) 2.7- 2.8 4.3- 4.7

22. Radiation Protection Everything placed into space must be protected in some way from cosmic radiation. Everything placed into space must be protected in some way from cosmic radiation. Typical commercial satellites protected to 2Mrad for a 10 year mission and SF of 2. Typical commercial satellites protected to 2Mrad for a 10 year mission and SF of 2. Our mission is shorter. Everything will be shielded to 1Mrad, unless otherwise required by specific components. Our mission is shorter. Everything will be shielded to 1Mrad, unless otherwise required by specific components. 3g/cm 2 of Al will provide an order of magnitude reduction in total dosage over a 10 year mission: this is sufficient to reduce total dosage to less than 1 Mrad. 3g/cm 2 of Al will provide an order of magnitude reduction in total dosage over a 10 year mission: this is sufficient to reduce total dosage to less than 1 Mrad. Sensitive components shall be organized as to benefit from spot- shielding. Sensitive components shall be organized as to benefit from spot- shielding. This layout must also mesh with the thermal management system. This layout must also mesh with the thermal management system.

23. Radiation Protection Reduction in Exposure when utilizing Al shielding Reduction in Exposure when utilizing Al shielding *Courtesy of http://see.msfc.nasa.gov/ire/iretech.htm

24. Power Requirements Requirements –Supply power to satellite –Support mission profile and requirements Major Design Drivers Major Design Drivers –Must fit inside Taurus launch vehicle –Provide adequate power –Reliable and easy to obtain

25. Power Trade Studies Trade Studies –Solar Power vs. RTG vs. Nuclear Reactor –Silicon vs. GaAs solar cells –Body Mounted Array vs. Gimbaled Array Panels

26. Power Triple Junction GaAs Solar Cells Triple Junction GaAs Solar Cells –28.5% average efficiency –26.6 square cm –Radiation Resistant –289 Watts per square meter Secondary Batteries Secondary Batteries Peak Power Trackers Peak Power Trackers –Active power regulation –Protects sensitive electrical components

27. Power Solar Array Peak Power Tracker Discharge Controller Loads Charge Controller Battery 28V Bus

28. Power - Battery Lithium Ion Lithium Ion –Highest number of Cycles for chosen depth of discharge (DOD) –High energy density –Customizable size and shape

29. Satellite Structure - Transport Fitting within the Taurus launch vehicle

30. Satellite Structure - Functional

31. Rest of Semester Detailed end-to-end data flow from lander and camera back to Earth including timeline, data rates, use of on-board storage, and contact times Detailed end-to-end data flow from lander and camera back to Earth including timeline, data rates, use of on-board storage, and contact times Optimize power system for our mission/components Optimize power system for our mission/components Detailed design of spacecraft structure and thermal conduction/radiation model Detailed design of spacecraft structure and thermal conduction/radiation model Integrate radiation protection with components and structure Integrate radiation protection with components and structure

32. Questions?