Shackleton Crater
Reconnaissance Mission
FDR
Trevor Fedie
Jason Breeggemann
Brian Evans
Mike Gavanda
Matt Gildner
Jeromie Hamann
Brian Nackerud
Andrew Smude
Jordan Stewart
Project Overview

Image 5 km annular region around Shackleton
Crater (at Moon’s south pole)
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–

3 color imagery
10 cm resolution
Image 80% of region in one month
Transmit to Earth once per day
Act as communication relay for lunar lander
exploring interior of crater
– Receive 2.5 Mbps S-band from lander
– Transmit to Earth once per day
Project Overview

Baseline orbit of 30 x 216 km
– Period approximately 120 minutes
– 12 orbits each day

Spacecraft must fit into Taurus Launch Vehicle
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–
–
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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
Major Tasks
Develop an operational concept to support
the mission objectives
 Determine the required data storage and
return-link data rate
 Determine the antenna size and
transmitter power
 Develop spacecraft design concept to
support the mission
 Demonstrate that spacecraft design meets
all mission requirements

Expectations

Spacecraft design will include
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–
–
–
–
Antenna size
Data storage size
Data link rates
RF link calculations
Mass budget
Power budget
Solar array size
Battery size
Pointing requirements
Stability requirements
Expectations

Operational concept will describe end-to-end
data flow from lander and from camera back to
Earth
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–
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
Timeline
Data rates
Usage of on-board storage
Contact times
Students will identify major problems and risks
associated with this mission concept
Mission Profile
Launch using Taurus Launch Vehicle
 Make burn for Earth-Moon transit
 De-spin
 Burn to enter Lunar orbit
 Begin imaging region around crater
 Finish imaging crater in about 14 days
 Lunar lander operations begin 4 months
after launch and last for one year

Camera
Crater Imaging Strategy
(Note: not to scale)
Camera

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
Storage Requirements

1.2 Gbits/photo
 Post compression
2 photos/pass
 12 passes/day
______________
 28.8 Gbits/per day


Considerations
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–
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Limit broadcasts
Safety margin
Earth Weather
Moments of “zero
visibility”
Data Storage
– SSR
 P9 family
 160 Gbits BOL
 Built-in FELICS
 256 Mb SDRAM
Computing & Data Storage
– 3U Compact PCI
 PowerPC RAD750
 Enhanced Power PCI Bridge
Communications

High Gain
– Cassegrain Antenna
 D=.98m
 d=0.3m
 Power usage 100 W
for broadcasting
 Double reflection
– Smaller dish
geometry
– Wiring
Earth Communications

Transmit high resolution photos
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Only one contact with White Sands per day
About 28.8 Gbits a day in pictures
Ka-band 26 GHz parabolic dish
Data rate up to: 30 Mbps
Power: 40 W
Gain: 44 db
White Sands receiving: 45 db/T
http://www.grc.nasa.gov/WWW/RT/2005/PT/PTN-wilson.html
Missions to the Moon
Chang’e 1 (CNSA)
 SMART-1 (ESA)
 Chandrayaan-1 (ISRO, India)

– X-band and S-band

KAGUYA (JAXA)
– S-band

LRO (NASA)
– KA-band
– 100 Mbps data rate
Link Analysis
SNR=Pr/Noise
 Pr=EIRP*Gr/losses
 Examples of losses

– Path loss (λ/4πR)2
– Atmospheric loss
– Polarization loss

N=k*T*B
Link Analysis (Earth Communications)
Data Rate
Frequency
Range
Path Loss
Power
Gain
Antenna Beam width
EIRP
Atmospheric Losses
Receiver Gain
Power Received
System Noise
SNR
Link Margin
Mbps
GHz
m
W
dB
deg
W
db
dB
W
W
dB
dB
db
74.77
103.98
85.85
-232.10
16.02
45.17
30
26
3.84E+08
6.17E-24
40
45.17
0.86
7.23E+05
58.59
-5.2
-5.20
45
45.00
7.42E-15 -141.30
4.14E-16
-153.83
17.91
12.53
10.03
Link Analysis (Rover Communications)
Parameters
Data Rate
Frequency
Range
Path Loss
Transmitter Power
Transmitter Gain
EIRP
Pointing loss
Gain
Power Received
System Noise
SNR
Link Margin
Mbps
GHz
m
W
dB
W
dB
dB
W
W
dB
dB
2.5
2.3
3.00E+05
1.20E-15
5
10
27.5
-3
10
1.03E-13
2.07E-14
3.48
dB
63.98
93.62
46.99
-149.22
6.99
10.00
14.39
-3.00
10.00
-131.42
-136.84
5.42
2.92
Rover Communications

Relay for rover in crater
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Required data rate of 2.5 Mbps
Required S-band from rover
2.3 GHz Omni-directional transmitter
Power: 5 W
Gain: 10 db
Rover Gain: 10 db
 L3 T&C transceiver MSX-765

Store data onboard satellite until downlink to White
Sands ground Station
http://www.l-3com.com/tw/telemetry_west/products/cxs_2000.htm
Commands
Receive commands from White Sands
 128 Kbps data rate
 Save commands and
relay to rover
 Required gain of 5 db
 Frequency of 2.1 GHz

http://lro.gsfc.nasa.gov/gallery-hardware.html
STK Model

Earth
– 4230 sec average daily window
– Tracking
 Azimuth: 0.01617 deg/s
 Elevation: 0.0154 deg/s

Moon
– 12 communication windows
 ~5 min over crater
 ~9 Gbits of data a day
 ~5 minutes required to send to Earth
STK Video
Spacecraft Data Flow
SSR
Camera
Transponders
Flight
Computer
Antennas
Spacecraft Data Flow

Crater Imagery
– Take 24 pictures per day (4048 x 60000 pixels)
– Images initially 14 bits/pixel
– Use look-up tables (LUT) to compress images to 8
bits/pixel before storage in camera memory
– Use fast efficient and lossless image compression
(FELICS) to compress images before storage in SSR
(FELICS compression >1.6:1)
– Transmit to White Sands once per day (~17 min
contact window required)
– Imaging operations complete within 14 days
Spacecraft Data Flow

Camera On-Board Memory Usage (1 July – 9 July 2009)
Spacecraft Data Flow
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Camera On-Board Memory Usage (zoomed in)
Spacecraft Data Flow
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Camera SSR Usage (1 July – 9 July 2009)
Spacecraft Data Flow

Rover Communications Relay
– Rover operations begin 4-5 months after launch of
first spacecraft
– Contact with rover 12 times per day (~60 mins total)
– Rover sends data at 2.5 Mbits/sec
– Collect approximately 9 Gbits of data each day
– Transmit to White Sands once per day (~5 min
contact window required)
– Rover operations complete within 17 months of
launch of first spacecraft
Spacecraft Data Flow

Rover Data SSR Usage (14 Dec – 22 Dec 2009)
Attitude Determination and Control

Derived Requirements
– Pointing stability (During an exposure time of 7.5e-3 sec)
-Roll axis <0.343 arcsec
-Pitch axis <0.343 arcsec
-Yaw axis <13.4 arcmin
– Pointing Accuracy with SF of 4
-Roll axis <2.5 arcmin
-Pitch axis <8.6 arcmin
-Yaw axis <2.7 arcmin
– Maneuverability
-3-axis control
-Pointing reassignment as fast as 90 deg in 6 minutes
Attitude Determination and Control

SED 16 Autonomous Star Tracker by Sodern
– Accuracy bias plus noise
-Roll axis <108 arcsec
-Pitch axis <36 arcsec
-Yaw axis <36 arcsec
– 10 Hz update
– 25x25 deg field of view
– Mass: 2.9 kg (with Baffle)
– Average power: 10.7 W
Attitude Determination and Control

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
Attitude Determination and Control

Momentum build up
– Disturbance Torques
- Gravity gradients
- Solar pressure
- Internal
– Slewing Maneuvers
- Once a day

HR14 Constellation Series Reaction Wheels by
Honeywell
– Max Reaction Torque 0.2 N-m
– Momentum Capacity 50 N-m-s
– Mass: 8.5 kg
Reaction Wheel Disturbances
• Force from jitters of a single reaction wheel in a local frame
Reaction Wheel Disturbances
• Moment from jitters of a single reaction wheel in a local frame
Disturbance Torques

External
– Solar ~1.68e-6 N-m
– Gravity Gradient ~3.04e-7 N-m

Internal
– Deployables ~2.22e-5 N-m
– At 50% max rpm of RWA’s (50 Hz) <0.76 N-m

Maximum resulting displacement of a single pixel
during camera exposure
– Roll Axis <0.008 arcsec
– Pitch Axis <0.008 arcsec
– Yaw Axis <0.008 arsec
Propulsion Trade Study
Looked at Clementine, Ranger, Luna and
Surveyor series Lunar orbiters
 Not much available info
 Studied fuel types, thrust ranges, and
system configurations
 Keep it simple

Propulsion System
Rocket Motor:
EADS Astrium S400-12
-MMH/MON-1
-420N thrust
-0.135 kg/sec
-Total V: 1021m/s
Propulsion System

Twelve 4N thrusters
– MMH/MON-1 Fuel/Oxidizer
– EADS Astrium S4

Two fuel tanks
– Composite structure
– 3600 psi rated
– Lincoln Composites
– 0.07 cubic meter
Determination of Thermal
Environment
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Moon surface temp
Altitude and attitude
Lunar view factor
Intense reflected IR
from lunar surface.
Thus objects placement
will important
If thermo environment
is compromised
anywhere and active
system will be used
Dynamic Thermo Model
Matlab-Simulink model
 Lumped sum approach
(craft skin, propellant
tank and electronic
system
 Craft Skin is subject to
radiation from space,
sun, internal components
and the moon
 System and Tank subject
to conduction, internal
power generation and
radiation with the skin.
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Orbit assumed to have
the craft always pointing
at lunar surface.
Craft skin Temp a
function of alt, attitude,
solar flux, (space/lunar)
view factors, internal
conduction and internal
radiation
System and tank temp
controlled by PID loop
Model analyzed under
max shade and max sun
light conditions
Sketch of Thermodynamic Model
Workspace
Testing of Thermodynamic Model
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Adjust craft surface emissivity and absorptivity.
Check against both models for zero heater
condition equilibrium
Varied internal surface properties for internal
radiation heat transfer
Used a steady state heating method to
determine max and min heating watts needed
Begin PID tuning of heating elements
Recheck model
Recheck craft properties & tolerances
Finalize thermo model
Results
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Surface exterior of 5.0 mil Aluminum FEP
Teflon®
Interior surface of black paint
Electronics casing of anodized metal
Tank heater of 60 watts electronic heater
requires 300 watts
Best to keep craft pointed at the lunar surface
Since radiation is primary means of heat
transfer, surface properties are the primary
means of passive thermal control
Radiation Environment
 Transient
Particles: Protons and Heavy
Ions of the Elements (GeV-TeV)
 Trapped Particles: Consists of Protons
and Heavy Ions (100MeV) and Electrons
(10MeV)
 Space Plasma: Major issue is surface
charging of spacecraft (0.1MeV)
Radiation Effects
Total Ionizing Dose (TID): Degradation from
Ionizing radiation, typically from primary protons
and electrons. Causes Timing skews, leakage,
and threshold shifts. Possible to use a shield to
absorb lower energy electrons and protons.
 Non-Ionizing Energy Loss (NIEL): Non-ionizing
displacement damage due to protons, electrons,
and neutrons. Damage is permanent. Shielding
may be effective based on the geometry and
technology of the equipment.
 Single Event Effects (SEE): Due to ionization by
a single charged particle. Typically high energy
protons and neutrons. Due to high Energy,
shielding is not effective.

Radiation Environment of Shackleton
Mission
Energy (eV) Flux (cm -2) % Total Flux
2.00E+06
4.00E+06
6.00E+06
8.50E+06
1.25E+07
1.75E+07
2.25E+07
2.75E+07
>30MeV
Total
(eV/cm 2) % Total Energy
2.95E+11
1.18E+11
7.20E+10
7.00E+10
6.90E+10
4.20E+10
2.71E+10
1.89E+10
37.68%
15.07%
9.20%
8.94%
8.81%
5.36%
3.46%
2.41%
5.9E+17
4.72E+17
4.32E+17
5.95E+17
8.625E+17
7.35E+17
6.1E+17
5.2E+17
4.95%
3.96%
3.62%
4.99%
7.24%
6.17%
5.12%
4.36%
7.10E+10
9.07%
7.1E+18
59.56%
100% 1.19E+19
100%
7.83E+11
-Spacecraft Incident Solar Proton Fluences for 3 Solar Active Years.
-Data from “Lunar Reconnaissance Orbiter Project” Goddard Space
Flight Center
Trade Study
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•
•
•
Material Thickness (cm)
Al
0.25
Ti
0.15
Cu
0.08
W
0.04
Pb
0.06
Teflon
0.31
Comparison of shielding materials and
corresponding thickness (cm).
Only considered particles less than 30 MeV
Averaged particle energy to 8.5 MeV/particle
Initial fluence data from “Lunar Reconnaissance
Orbiter Project” Goddard Space Flight Center
Radiation Shielding
Based on LRO requirements, the
Shackleton Mission will be shielded to 10
krad total dosage rate.
 Utilize 0.25 cm Al shielding. Equipment
with Linear Threshold Limit less than
75MeVcm2/mg for destructive events and
37 MeVcm2/mg for non-destructive events
are adequately protected. Other
equipment will need further analysis.

Alternative/Additional Actions to Ensure
Survivability of Spacecraft
1.
Use of NMOS and PMOS, rather than CMOS increases
radiation resistance at the expense of power
consumption.
2.
Transistor-Transistor Logic (TTL) is more resilient but
uses more power.
3.
Independent Processors with “voting” logic, and extra
bits to allow error detection.
4.
Spot-Shielding of specially sensitive equipment
Power

Requirements
– Supply power to satellite
– Support mission profile and requirements

Major Design Drivers
– Must fit inside Taurus launch vehicle
– Provide adequate power
– Reliable and easy to obtain
Power

Trade Studies
– Solar Power vs. RTG vs. Nuclear Reactor
– Silicon vs. GaAs solar cells
– Body Mounted Array vs. Gimbaled Array
Panels
Power System Trade Study
Power System
Solar Panels
RTG
Nuclear Reactor
Weight
Low?
High?
High?
Size
Compact for launch, Compact for launch,
large in use
not large in use
?
Orientation Issues
Yes
Yes
Yes
Heat Generation
Low
High
High-Very High
Optimal Temp.
Yes
No
Yes
Optimal Light
Yes
No
No
Radiation Degradation
Yes
Yes
Yes
Affected by other
components
Yes
No
?
Affects other components
Some
Yes
Yes
Require Batteries
Yes
Maybe not
Maybe not
Reliability
Decent
High
Low-High
Availability
High
Medium
Low
Ease of installation
High
Very low
Low
Power Level
Med-High
Med-High
High-Very High
Longevity
7-10 years
7-10 years
7-10+ years
Power Budget
System
Max Power (W)
Avg. Power (W)
Voltage (V)
SED 16
10
28
SED 16
10
28
Reaction Wheels
195
22
28
Reaction Wheels
195
22
28
Reaction Wheels
195
22
28
Reaction Wheels
195
22
28
38
28
60
28
37
28
SIRU (gyros)
HiRISE Camera
125
SSR
S-Band System
35
8.5
28
Ka-Band System
110
8.5
28
35
35
28
400
200
28
1580
495
Rocket Motor
Heating
Totals
Mission Profile

Power Requirements
– Imaging: 560W
– Transmitting to Earth: 597W
– Transmitting to Rover: 522W
– Receiving (from Earth or Rover): 495W
– Orienting (Sun Side): 600W
– Orienting (Dark Side): 800W
Photovoltaic Cells

Basic Triple Junction GaAs Solar Cells
– 28.5% average efficiency
– 26.6 square cm
– Radiation Resistant
– ~396 W/m^2

800 Watts Max Power  1020 Cells
– 15 cells per 28V  4 15x8 cell panels and 4
15x9 cell panels
Secondary Battery

Lithium Polymer (Li-Po)
– High Energy Density
– Proven by use on other missions (MRO, LRO,
etc.)

28 Volt Bus  9 3.4V Cells
Power Regulation

Peak Power Tracking
– Active Power Regulation
– Protects Sensitive Electronics (Camera)
– Requires DC/DC Conversion to preserve
efficiency
– ~3kg in mass  small penalty
Power
Solar Array
Battery
DC/DC
Converter
Peak Power
Tracker
Charge
Controller
28V Bus
Loads
Discharge
Controller
Placement and Structure
Components
Satellite Structure - Functional
Taurus Launch Vehicle
Center of Gravity
Side View
Top View
Δy = 2.2 cm
Δz = 0.3 cm
Mass Budget
Component
kg Component
kg
Camera
65
Star Tracker
5
Computer
8
Main Thruster
7
SSR (Memory)
12
Guiding Thrusters
4
Battery
9
Propellant Tanks
25
Power tracker
3
5 mm Al FEP Teflon®
3
Solar Array
9
Blankets
5
Transponder
5
Heaters
3
Ka-band Transmitter
4
Harness
10
High-Gain Antenna
4
Structure
90
Reaction Wheels
34
Propellant Mass
121
Gyro
7
Total Mass
433
Objective for
everyday launch
opportunity:
445 kg
Mass Margin:
12 kg
Major Problems and Risks
Some redundancy built in – certain
equipment failures could disable
spacecraft
 Space debris impact could disable
spacecraft
 Single event (radiation) effect could
disable spacecraft

Questions?
Thanks to:
David Everett – NASA GSFC
Rajapandiyan Asaithambi
Professor Flaten
Professor Garrard
Professor Hammer
Also thanks to:
Candice Hansen – NASA JPL
Jim Bergstrom – Ball Aerospace