Shackleton Crater Shackleton Crater Reconnaissance Mission Reconnaissance Mission
PDRPDRTrevor FedieJason BreeggemannBrian EvansMike GavandaMatt GildnerJeromie HamannBrian NackerudAndrew SmudeJordan Stewart
Project OverviewProject Overview
Image 5 km annular region around Image 5 km annular region around Shackleton Crater (at Moon’s south Shackleton Crater (at Moon’s south pole)pole)– 3 color imagery3 color imagery– 10 cm resolution10 cm resolution– Image 80% of region in one monthImage 80% of region in one month
Act as communication relay for lunar Act as communication relay for lunar lander exploring interior of craterlander exploring interior of crater– 2.5 Mbps S-band from lander2.5 Mbps S-band from lander– Transmit to Earth once per dayTransmit to Earth once per day
Project OverviewProject Overview Baseline orbit of 30 x 216 kmBaseline orbit of 30 x 216 km
– Period approximately 120 minutesPeriod approximately 120 minutes– 12 orbits each day12 orbits each day
Spacecraft must fit into Taurus Launch VehicleSpacecraft must fit into Taurus Launch Vehicle– Goal of 445 kg for everyday launch opportunitiesGoal of 445 kg for everyday launch opportunities– Must handle spin stabilized upper stage (60 rpm)Must handle spin stabilized upper stage (60 rpm)– Must interface with launch vehicle upper stageMust interface with launch vehicle upper stage– Must fit into launch vehicle envelopeMust fit into launch vehicle envelope
Mission ProfileMission Profile
Launch using Taurus Launch VehicleLaunch using Taurus Launch Vehicle De-spin, make burn for Earth-Moon De-spin, make burn for Earth-Moon
transittransit Burn to enter Lunar orbitBurn to enter Lunar orbit Begin imaging region around craterBegin imaging region around crater Send data back to Earth once per daySend data back to Earth once per day Finish imaging crater in about 14 daysFinish imaging crater in about 14 days Lunar lander operations begin 4 months Lunar lander operations begin 4 months
after launch and last for one yearafter launch and last for one year
Spacecraft Data FlowSpacecraft Data Flow
Flight Compute
r
SSR
Transponders
Star Tracker
Camera
Gyroscope
Reaction Wheels
Rocket Motor
Thrusters
Temp Sensors
Antennas
CameraCamera
Crater Imaging Strategy (Note: not to scale)
CameraCamera
HiRISE (used on Mars HiRISE (used on Mars Reconnaissance Orbiter)Reconnaissance Orbiter)– Pushbroom TDI imager (4,048 pixels Pushbroom TDI imager (4,048 pixels
across swath)across swath)– 0.5 m aperture0.5 m aperture– 28 Gbits internal data storage28 Gbits internal data storage– Internal LUT image compressionInternal LUT image compression– Mass: 65 kgMass: 65 kg– Average power: 60 WAverage power: 60 W
Computing & Data StorageComputing & Data Storage
– 3U Compact PCI3U Compact PCI PowerPC RAD750PowerPC RAD750 Enhanced Power PCI BridgeEnhanced Power PCI Bridge
– SSRSSR P9 familyP9 family 160 Gbits BOL160 Gbits BOL Built-in FELICSBuilt-in FELICS 256 Mb SDRAM256 Mb SDRAM
CommunicationsCommunications
Earth Comm.Earth Comm.– Cassegrain AntennaCassegrain Antenna
D=.98mD=.98m d=0.3md=0.3m Power usage 40 WPower usage 40 W Double reflectionDouble reflection
Earth CommunicationsEarth Communications
Transmit high resolution photosTransmit high resolution photos– Only one contact with White Sands per dayOnly one contact with White Sands per day– Around 90 Gbits a day in picturesAround 90 Gbits a day in pictures– Ka-band 26 Ghz parabolic dishKa-band 26 Ghz parabolic dish– Data rate:100 MbpsData rate:100 Mbps– Power: 40 WPower: 40 W– Gain: 44 dbGain: 44 db– White Sands receiving: 45 db/TWhite Sands receiving: 45 db/T
Similar to LROSimilar to LRO
Rover CommunicationsRover Communications
Relay for rover in craterRelay for rover in crater– Required data rate of 2.5 MbsRequired data rate of 2.5 Mbs– Required S-band from roverRequired S-band from rover– 2.3 Ghz Omni-directional transmitter2.3 Ghz Omni-directional transmitter– Power: 5 WPower: 5 W– Gain: 5 dbGain: 5 db– Rover Gain: 5 dbRover Gain: 5 db
L3 T&C transceiver MSX-765L3 T&C transceiver MSX-765
Store Data onboard satellite until downlink to Store Data onboard satellite until downlink to White Sands ground StationWhite Sands ground Station
CommunicationsCommunications
EarthEarth– 4230 sec average daily window4230 sec average daily window– TrackingTracking
Azimuth: 0.01617 deg/sAzimuth: 0.01617 deg/s Elevation: 0.0154 deg/sElevation: 0.0154 deg/s
MoonMoon– 12 communication windows 12 communication windows
6.533 min over crater6.533 min over crater 117.6 Gbits of data a day117.6 Gbits of data a day 2 min required to send to earth2 min required to send to earth
Pointing stability/accuracyPointing stability/accuracy• • Torque disturbances must result in 10 cm Torque disturbances must result in 10 cm
or less or less ground displacements during ground displacements during exposure timeexposure time
• • Attitude accuracy Attitude accuracy
-Roll axis <2.5 arcmin-Roll axis <2.5 arcmin
-Pitch axis <8.6 arcmin-Pitch axis <8.6 arcmin
-Yaw axis <2.7 arcmin-Yaw axis <2.7 arcmin
ManeuverabilityManeuverability• • 3-axis control3-axis control• • Pointing reassignment as fast as 90 deg Pointing reassignment as fast as 90 deg
in 6 in 6 minutes minutes
Attitude Determination and Attitude Determination and ControlControl
Attitude Determination and Attitude Determination and ControlControl
SED 16 Autonomous Star Tracker by SED 16 Autonomous Star Tracker by SodernSodern– 36 arcsec in pitch/yaw, 108 arcsec roll 36 arcsec in pitch/yaw, 108 arcsec roll
(bias plus noise)(bias plus noise)– 10 Hz update 10 Hz update – 25x25 deg field of view25x25 deg field of view– Mass: 2.9 kg (with Baffle) Mass: 2.9 kg (with Baffle) – Average power: 10.7 WAverage power: 10.7 W
Attitude Attitude DeterminationDetermination and and ControlControl
Scalable SIRU (gyro) by Northrop Scalable SIRU (gyro) by Northrop Grumman Grumman – Achieves Gyro Bias stability of 0.0003 deg/hrAchieves 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 noiseLow noise– Mass: 7.1 kgMass: 7.1 kg– Average power: 38 WAverage power: 38 W
Attitude Determination and Attitude Determination and ControlControl Momentum build upMomentum build up
•• Disturbance TorquesDisturbance Torques - Gravity gradients- Gravity gradients- Solar pressure- Solar pressure- Internal- Internal- Deployables- Deployables
•• Slewing Maneuvers Slewing Maneuvers - Minimum twice- Minimum twice a day a day
HR14 Constellation Series Reaction HR14 Constellation Series Reaction Wheels by Honeywell Wheels by Honeywell – Max Reaction Torque 0.2 N-mMax Reaction Torque 0.2 N-m– Momentum Capacity 50 N-m-sMomentum Capacity 50 N-m-s– Mass: 8.5 kgMass: 8.5 kg
Propulsion SystemPropulsion SystemRocket Motor:Rocket Motor:
-EADS Astrium S400-12 -EADS Astrium S400-12
-MMH/MON-1 -MMH/MON-1
-420N Isp:318s-420N Isp:318s
-Total -Total V needed: 1011m/sV needed: 1011m/s
Propulsion SystemPropulsion System
Twelve 4N thrustersTwelve 4N thrusters– MMH/MON-1 Fuel/OxidizerMMH/MON-1 Fuel/Oxidizer– EADS Astrium S4EADS Astrium S4
Two 0.05 cubic meter tanksTwo 0.05 cubic meter tanks– Composite structureComposite structure– 3600 psi rated3600 psi rated– Lincoln CompositesLincoln Composites
Determination of Thermal Determination of Thermal EnvironmentEnvironment
Moon surface temp Moon surface temp Altitude and attitudeAltitude and attitude Lunar view factorLunar view factor Intense reflected IR Intense reflected IR
from lunar surface. from lunar surface. Thus objects Thus objects
placement will placement will importantimportant
If thermo enviroment If thermo enviroment is compromised is compromised anywhere and active anywhere and active system will be usedsystem will be used
Thermal AnalysisThermal Analysis Equilibrium temperature range of 270-310 Equilibrium temperature range of 270-310
KelvinKelvin Electric heaters and sensors on items that do Electric heaters and sensors on items that do
not fall within their ideal ranges.not fall within their ideal ranges. Use of heat tubes and radiators if neededUse of heat tubes and radiators if needed Will be tuned in by a dynamic modelWill be tuned in by a dynamic model
Structural System Structural System DeterminationDetermination
Properties of Ti an AlProperties of Ti an Al Bulk system of TI. Bulk system of TI. AL only used where AL only used where
conduction requires conduction requires it.it.
Current structural Current structural mass of mass of approximately 18Kg.approximately 18Kg.
Titanium will also Titanium will also have less thermal have less thermal expansionexpansion
ALAL TiTi
Yield Yield Strength Strength (MPa)(MPa)
475-475-520520
480-480-11711700
Coefficient of Coefficient of Thermal Thermal expansion expansion
≈≈2323 8-118-11
ConductivitConductivityy
(W/m - K)(W/m - K)
88-88-210210
6-176-17
DensityDensity
(Mg/m³)(Mg/m³)2.7-2.7-2.82.8
4.3-4.3-4.74.7
Radiation ProtectionRadiation Protection Everything placed into space must be protected in some Everything placed into space must be protected in some
way from cosmic radiation.way from cosmic radiation. Typical commercial satellites protected to 2Mrad for a 10 Typical commercial satellites protected to 2Mrad for a 10
year mission and SF of 2.year mission and SF of 2. Our mission is shorter. Everything will be shielded to Our mission is shorter. Everything will be shielded to
1Mrad, unless otherwise required by specific components.1Mrad, unless otherwise required by specific components. 3g/cm3g/cm22 of Al will provide an order of magnitude reduction in of Al will provide an order of magnitude reduction in
total dosage over a 10 year mission: this is sufficient to total dosage over a 10 year mission: this is sufficient to reduce total dosage to less than 1 Mrad.reduce total dosage to less than 1 Mrad.
Sensitive components shall be organized as to benefit from Sensitive components shall be organized as to benefit from spot-shielding.spot-shielding.
This layout must also mesh with the thermal management This layout must also mesh with the thermal management system.system.
Radiation ProtectionRadiation Protection
Reduction in Exposure when utilizing Al shieldingReduction in Exposure when utilizing Al shielding
*Courtesy of http://see.msfc.nasa.gov/ire/iretech.htm*Courtesy of http://see.msfc.nasa.gov/ire/iretech.htm
PowerPower
RequirementsRequirements– Supply power to satelliteSupply power to satellite– Support mission profile and requirementsSupport mission profile and requirements
Major Design DriversMajor Design Drivers– Must fit inside Taurus launch vehicleMust fit inside Taurus launch vehicle– Provide adequate powerProvide adequate power– Reliable and easy to obtainReliable and easy to obtain
PowerPower
Trade StudiesTrade Studies– Solar Power vs. RTG vs. Nuclear ReactorSolar Power vs. RTG vs. Nuclear Reactor– Silicon vs. GaAs solar cells Silicon vs. GaAs solar cells – Body Mounted Array vs. Gimbaled Array Body Mounted Array vs. Gimbaled Array
PanelsPanels
PowerPower
Triple Junction GaAs Solar CellsTriple Junction GaAs Solar Cells– 28.5% average efficiency28.5% average efficiency– 26.6 square cm26.6 square cm– Radiation ResistantRadiation Resistant– 289 Watts per square meter289 Watts per square meter
Secondary BatteriesSecondary Batteries Peak Power TrackersPeak Power Trackers
– Active power regulationActive power regulation– Protects sensitive electrical componentsProtects sensitive electrical components
PowerPower
Solar Array
Peak Power
Tracker
Discharge Controller
Loads
Charge Controller
Battery
28V Bus
Power - BatteryPower - Battery
Lithium IonLithium Ion– Highest number of Cycles for chosen Highest number of Cycles for chosen
depth of discharge (DOD)depth of discharge (DOD)– High energy densityHigh energy density– Customizable size and shapeCustomizable size and shape
Satellite Structure - Satellite Structure - TransportTransport
Fitting within the Taurus launch vehicle
Satellite Structure - Satellite Structure - FunctionalFunctional
Rest of SemesterRest of Semester
Detailed end-to-end data flow from lander Detailed end-to-end data flow from lander and camera back to Earth including and camera back to Earth including timeline, data rates, use of on-board timeline, data rates, use of on-board storage, and contact timesstorage, and contact times
Optimize power system for our Optimize power system for our mission/componentsmission/components
Detailed design of spacecraft structure Detailed design of spacecraft structure and thermal conduction/radiation modeland thermal conduction/radiation model
Integrate radiation protection with Integrate radiation protection with components and structurecomponents and structure
Questions?Questions?
