Lunette: Satellite to Satellite
Gravity Mapping of the Moon
Maria Short
October 24, 2007
9th ILEWG International Conference on Exploration and Utilisation of the Moon
Authors:
M. Short, C. Short, A. Philip, J. Gryzmisch,
R. Zee, H. Spencer, and J. Arkani-Hamed
Space Flight Laboratory
Who are We?
•
Unique university lab in Canada focusing on
microspace systems research and development
“Microspace” = disciplined small team approach to
using the latest commercial technologies in space
October 24, 2007
•
Developed key subsystems for MOST
and supported integration, test and operations
•
Canadian Advanced Nanospace eXperiment (CanX)
nanosatellite program to train students and provide
for low cost space access
•
Technology research in communications, propulsion,
radiation testing
•
Full-time professional staff with microspace systems
expertise
•
Facilities to support the development and
qualification of space systems
Lunette: Satellite
to Satellite Gravity
Mapping of the Moon
www.utias
-sfl.net
2
Introduction
• Farside gravity map of the Moon is not as good as the nearside map;
Gravity map is important for navigation and exploration
• Lunette is a mission concept involving a nanosatellite in formation with a
parent satellite around the Moon. Whole sphere maps are possible with
radio range-rate measurements between the satellites
• UTIAS/SFL has developed the Generic Nanosatellite Bus (GNB) in support
of BRITE Constellation (CanX-3) and the CanX-4/CanX-5 formation flying
mission
• UTIAS/SFL has customizable separation systems, “XPODs,” that can be
used to eject GNB satellites from launch vehicles or parent satellites
• GNB and XPOD technology can be used to support the Lunette mission.
Lunette is a “portable” nanosatellite mission
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Lunette: Satellite to Satellite Gravity Mapping of the Moon
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Mission Concept
• Science mission
– To map Lunar farside gravity field, to
10-20 mGal gravity surface anomaly
• Free-flying nanosatellite, ejected from and
flying in formation with a parent satellite,
both in low Lunar orbit, measuring relative
range rate using radio tracking
• Science instrument
– Ranging radio transponder
• Bus needs 3-axis attitude control and
propulsion
• Phase A Study for European Student Moon
Orbiter (ESMO) under ESA’s SSETI Program.
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Analysis Method
• The range rate signal is approximately
proportional to selenopotential difference along
the track of the two satellites
• The proportionality constant is the orbital
velocity.
U12 ≅ Δv ⋅ vorbit
Where U12 is the potential difference between the locations
of the two satellites, Δv is the measured range-rate and v is
the orbital velocity
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Range Rate Accuracy at τ = 10s (mm/s)
Local Sensitivity Model
Distance from Leading Satellite to Mascon (km)
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Tracking Approach
Range Rate Measurements
• The range-rate tracking algorithm performs accurate
measurements of the relative speed of Lunette with
respect to ESMO
• Range-rate is derived by measuring the Doppler shift
caused by relative motion on a carrier transmitted by
ESMO and in turn sent back by Lunette.
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Tracking Approach
Range Rate Measurements
⇒ fD =
October 24, 2007
f1Δv ⎡ Δv ⎤
2+ ⎥
Mc ⎢⎣
c ⎦
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Communications
• Goal of communications design
– Perform high precision range-rate
measurements
– Perform low precision range
measurements
– Low-speed two-way data link
between Lunette and ESMO
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Propulsion
• Propulsion required to maintain
circular orbit and perform 1°
plane change maneuver
• Estimate about one thrust per
week is required to maintain
orbit
• Plane change maneuver is
about half the available ∆V
• Using four thrusters propulsion
system is beneficial for
momentum dumping of the
three axis attitude control
system
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Perilune Attitude (km)
Perilune Altitude Evolution
Time (days)
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Attitude Determination and
Control System
Attitude goals during science measurements:
• Null rotation rates to prevent data corruption
• Coarse Pointing of antennas
Achieved with:
• Rate Sensors
• Sun Sensors
• Occasional Star tracker usage
• Reaction Wheels
Expected effect of ADCS errors
on range rate measurement: 0.038mm/s
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On-Board Computers
• The OBC baseline will be identical for Analyzer and
Lunette
– Simpler integration and workload
• Centralized Computer Architecture
• Uses three computers
– Two on Lunette sub-satellite
– One on Analyzer
• Must be able to store data for up to 15 hours
– Only transfer data to ESMO over poles twice a day
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Power System
• Power Generation with solar cells
– 27.5% efficiency
• Power Storage with Li-ion batteries
– 20 Wh energy capacity
• Direct Energy transfer
– Always within 5% of peak power at operating point for
mission life
• To increase power generation, satellite will be edge
pointing to the sun
• Battery Depth of Discharge always kept above 50%
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Structural and Thermal
Structural
• 25x25x25 cm3
• Layout dictated by orthogonal
antenna/star-tracker placement
and large fuel tank
Thermal
• Cannot survive prolonged lunar
eclipse (limits life to ~=
6months)
• Must survive both a ‘dawn-dusk’
orbit and a ‘noon-midnight’ orbit
due to lack of precession and 4
week lunar day
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Separation System
• Need to protect Lunette from
radiation dictated design
• Radiation shielding on board the
sub-satellite would increase weight
precluding orbital maintenance
maneuvers, therefore satellite is to
be encapsulated within ESMO
• Separation system design based
on SFL’s XPOD deployment system
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Conclusion
• 20mGal accuracy is feasible with
microsatellite/nanosatellite technology
• ESMO and Lunette is an ideal pairing of student projects
• Completed a Phase A study for ESMO. Subsequent
phases subject to funding and approval by ESA and CSA
• The challenges in measuring the lunar gravity field are
not primarily overcoming technical hurdles, but instead
simply the opportunity to achieve this goal
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