1. No category

Comet Sample Return
CSR
Design Review
March 20, 2003
University of California, Santa Cruz
Electrical Engineering 127 & 128
Acknowledgements
Alec Gallimore
Ken Lande
Greg Laughlin
Douglas Lin
Claire Max
Graham Smith
Martin Yabroff
Mission Statement
The Objective of the Comet Sample Return
project is to develop a viable, robust space
system for returning pristine comet nucleus
samples for analysis on Earth, thus enabling a
better understanding of the formation and
evolution of the Solar System and the origins
of life on Earth.
Why Sample a Comet?
• Comets are thought to
be the oldest remains of
the early Solar System
and are composed of the
same matter as the early
solar nebula.
• Comets are also thought
to be the main source
for water and organic
molecules on earth
Why Sample a Comet?
• National Research
Council
– “Only a returned sample
will allow the necessary
elemental, isotopic,
organic, and mineralogical
measurements to be
preformed.”
• Better understanding of
how the Solar System
formed and life on Earth
formed.
Ways to Understand the Early Solar System
Study
Pros
Cons
Geology
Comets
Asteroids
Computer
Modeling
Meteorites
•Human on
site analysis
•Earth based
•Pristine
•Easier to
samples from land on than
the eerily
comets
solar system
•Easier to drill
than asteroids
•Easy to do
•Earth
based
•Extraterres
trial Sample
•Earth
based
•The earth is
constantly
changing over
time
•Difficult to
land on
•Run
present state
backwards
•Numerous
outcomes
•No initial
conditions
•Sample
can change
chemically
while
entering the
atmosphere
•Difficult to
drill into
Scientific Objectives of CSR
• To study:
– Internal Structure
– Ion and Dust tail
– Organic Parent
Molecules
– Rotation
– Activity while
Approaching the Sun
• To Return:
– Pristine Samples of a
Comets Nucleus
– Pictorial Atlas of the
Surface of a comet
– Sensor Data
•
•
•
•
•
Rotational
Spectrometer
Permittivity
Magnetometer
Dust Analyzer
Comet Exploration
• Giotto, Suisei, Sakigake,
Vega 1 & 2 (1985) Halley
flyby
• ICE (1985) Comet GiacobiniZinner flyby
• Stardust (2004) Dust sample
collection and return
• Rosetta (Planned 2003
launch) Orbiter and Landers
• Deep Impact (2004 launch)
Sensor probe to impact
comets surface
• CSR the next step
Mission Overview
Mission Overview
•
Rendezvous and trail a comet, at a
distance of approximately 1-1.7 AU
from the sun, entering into orbit
around the comet around 1.7 AU.
•
Provide ~30 days mapping of comet,
with the resolution needed to select
an appropriate landing site.
– Full surface mapping 10x
– Potential landing site mapping 50x
• 5-10 sites each mapped 5-10
times
Encke Orbit
Mission Overview (Cont)
• Land on the comet at
approximately 2 AU from the Sun
and provide real time data on
comet environment during the
landing process
• Deploy Smart Dust seismic motes
on the surface of the comet for
seismic sounding
• Drill at least 0.3m deep to obtain a
sample of the comet
• Perform a preliminary in-situ
compositional analysis of the
sample
Mission Overview (Cont)
• Deploy coma and tail probes to orbit the comet
while it approaches the sun
• Maintain the sample at the comets ambient
temperature (50-200° K) for chemical preservation
during voyage back to Earth.
• Final result: Return at least 100g of preserved comet
material to a Lower Earth Orbit to be recovered by
the Space Shuttle.
Candidate Comets
Wild 2
Encke
Encke
Orbiter and Lander Configuration
Orbiter & Lander
Configuration
Return
Configuration
Lander
after separation
Lander Configuration
during Drilling Ops
Innovations
• Wireless networked
miniature sensors (Smart
Dust, Micro Motes)
• Seismic imaging of the
interior of the comet
• Portable drilling and
sample collection
technology
• On site sample analysis
devices
Project Overview: Teams
UCSC and University of Michigan Joint Effort
UCSC
U of M
• Mission Science &
Architecture
• Structures & Materials
• Nucleus Sample
Return & Analysis
• Power & Thermal
• Remote Sensing &
Analysis
• Smart Dust
• Communications
• Payload Computers
Complete
System
Design
• Trajectory & Propulsion
• Guidance, Navigation,
Command, Control, &
Communications
• Test & Quality Control
• Budget & Scheduling
Joint Tasks:
Mission Science
Communications
Michigan Preliminary Study
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
Mission Science and Architecture
Deutron Kebebew
Jason Canoy
Outline
•
•
•
•
•
Main Objective
Background
Mission Architecture
Mission Science
Mission Conclusion
www. comet-cartoons. com
Main Objective
• To provide humanity with a better explanation
and understanding of the source of organic and
non-organic molecules and their relation to
Comets and life on Earth.
• To give insight on the evolution of comets
from their origins to their destiny, and to
understand the source and evolution of the
solar system, the planets, asteroids, and space
debris.
Why Comets?
Determine
source
space debris
Start
Determine comets
composition
Understand
how nebulas
were formed
Understand
how life
appeared on
Earth
Understand how the Solar
System was formed
What Are Comets?
• Comets are small, fragile,
irregularly shaped bodies
composed of mixture of
grains and frozen gases.
• Comet structures are diverse
and very dynamic, but they
all develop a surrounding
cloud of diffused material.
• As comets approach the Sun
they develop enormous tails
of luminous material that
extends for millions of
kilometers from the head.
Life Cycle of a Comet
6AU, comet inactive
4AU, comet becomes active
Tail can be
1-2AU long
Comet has welldeveloped tail
Earth’s Orbit
Comet at
perihelion
6AU, comet inactive
again
Jupiter’s Orbit
Anatomy of a Comet
Nucleus
1-50 km
Bluish ion tail
106-107 km
Coma
104-105 km
Yellowish dust tail
106-107 km
Target Trade Study
Best =
Comet
Orbital
Period
Perihelion
Distance
Semi-Major
Axis
Orbital
Inclination
Encke
3.3 yrs
0.340 AU
2.21 AU
11.8 deg
Wild2
6.39 yrs
1.583 AU
3.44 AU
3.2 deg
Subject Profile
• Comet P2/Encke
Advantages:
– It provides clear understanding for the source
for dust particles in the inner solar system
– It provides insight to the evolution of comets,
the solar system and asteroids
– Short term space mission and cost effective
due to Encke’s short orbital period
– It provides a layer highly organic rich surface
due to the many encounters with solar
wind and solar light
– Possibly rock like structure for the nucleus
easy for landing and less gases activity
Disadvantage:
– It provide less of pristine volatile material
that is found on comets
Mission Architecture
• Spacecraft will consist of:
–
–
–
–
Orbiter
Lander
Coma Probe
Tail Probe
• Spacecraft will rendezvous with
comet at 1.7-1.9AU.
• Orbiter will hold Lander and
probes until deployment.
Orbital Relation of Comet Encke
Deploy tail and coma
probes receive and
analysis data
Continue monitoring the
activities of the comet
0.34 AU
Earth orbit
1AU
Mars orbit
1.5AU
Lander live the comet
surface
Encke orbit
2.2 AU
Jupiter orbit
5.2 AU
Start following the comet
Rendezvous with comet
and start imaging,
mapping and remote
sensing
Deploy the Lander and
start extraction of sample
The Orbiter
• Acts as the super node
– Communicates with Earth,
Lander, and probe
• Orbiter Probe measurements
–
–
–
–
–
Surface and Nucleus imaging
IR and UV Spectra
Ion mass measurements
Surface mapping
Magnetic field
spaceprojects.arc.nasa.gov/ .../pioneer/ PNhist.html
The Lander
• Sample extraction
• Onboard sample analysis
– Analysis of chemical, mineral,
biological, and isotopic
compositions.
– Near surface strength
– Density and texture
– Thermal measurement
• Surface imaging
• Seismic signal analysis
• Return sample to Earth
www.mssl.ucl.ac.uk/www_plasma/ missions/rosetta.html
Coma Probe
• Coma Probe measurements
– Measure the material production
rate
– Measure the different isotope CO
and CO2
– Monitor the evolution and
production CO, CO2, and H2O
– Rational energy of the partials
– Photolytic process
– Physicochemical process (formation)
– Measure neutral gas
• Hydrogen Envelope
– Measure hydrogen isotopes
www.mssl.ucl.ac.uk/www_plasma/ missions/rosetta.html
The Tail Probe
• Tail Probe measurements
– Dust tail analysis
• Composition dust particles
• Partial size, density, and velocity
• Optical prosperities of dust
particles
– Ionic tail analysis
•
•
•
•
Ion pressure
Magnetic field
Ionic particles
Plasma
http://www.boeing.com/defense-space/space/bss/factsheets/scientific/pioneer/pioneer.html
Orbiter
Coma
probe
Lander
Tail
probe
Mission Architecture
Architecture
Cost
Weight
Scientific
Benefits
Mission
~180W
≈ 100kg
Sample
surface and
core of
nucleus
Mars
observer
Analysis of
coma and
nucleus
Rosetta
Power
Lander
≈ 100150M
Orbiter/Lander
≈ 170200M
~270W
≈ 250290kg
Orbiter/Probe
≈ 150175M
~120W
≈ 200270kg
Analysis of
coma and tail
Star Dust
Lander/Probe
≈ 100150M
~180W
≈ 200210kg
Analysis of
nucleus and
tail
Pioneer
Lander/Orbiter/
Probes
≈ 170250M
~300W
≈ 420470kg
Analysis of
entire comet
CSR
Mission Science Objectives
• Determine characteristics of nucleus
– Mass, size, composition
• Study comet environment
• Analyze organic compounds and isotopes
– Are these molecules connected with the origin of
life?
Mission Science
• Remote Measurements
–
–
–
–
–
–
–
–
–
Imaging
Mapping
Spectrometer readings
Rotation
Magnetic fields
Pressure
Gas expulsion rates
Particle velocities
Photolytic
physicochemical process
• In Situ Measurements
–
–
–
–
–
–
–
Compound abundance
Particle velocities
Density of nucleus
Nucleus composition
Near surface strength
Porosity
Seismic
Important Measurements
Primary =
Secondary =
What
How
Why
Mass/Size
Radio Science Investigation
(Rosetta)
Obtain better understanding
Near Surface Strength
Sonic Drill
of comet nucleus.
Isotopes
UV Spectrometer
Determine connection with
C, N, H, O
IR Spectrometer
comets to origin of life.
EM Fields
Magnetometer
Radio Noise
Heterodyne Radio Receiver
Temperature
Smart Dust
Determine characteristics of comet
Total Pressure
Smart Dust
Determine characteristics of comet
Determine characteristics of comet
Determine gas production rate of
comet
Conclusion
• Encke is the best target for our mission.
• Design an Orbiter/Lander/Probes to carry out
the mission.
• Desire measurements that describe the nucleus
and indicate origins of life.
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
Lander Sensors and Sample Collection
Kaeton Lee
Adam Seger
Austin Roman
Outline
• Drilling
• Seismic sources
• Sample encapsulation
and cooling
• In situ analysis
Requirements
• Drilling
–
–
–
–
100g samples
Depth greater than 0.3m
Pristine sample
Power/Mass/Volume
• Sample Storage
– Find sample temperature
– Keep sample at temp. for
entire trip back.
– Power/Mass/Volume
• In-situ Measurements
– Obtain scientific analysis
incase returned samples are
not useable or pristine.
– Power/Mass/Volume
• Seismic source
– Produce seismic source for
smart dust analysis
– Power/Mass/Volume
What Drill to Use?
SATM
(sample acquisition and transfer mechanism)
vs.
USDC
(ultrasonic/sonic drill corer)
SATM Features
-Acquire samples to 1.2
meters below surface
-Features in-situ
instruments for analysis
and return containers
-Developed for low
temperature operation
-Ability to prevent cross
contamination
USDC features
- Small light-weight
mechanism of drilling and
coring with integrated sensors
- Can be used as a seismic
source for mapping comet
- No drill bit sharpening
- Drilling technique prevents
clogging during core sampling
- Ability to take in-situ analysis
and core samples for return
USDC vs. SATM
DRILL
POWER
(W)
axial
PreLoad
(N)
MASS
(Kg)
Sample
Size/
depth
of
sample
Drilling
Speed
Maintenance
Integrated
Sensors
Autonomous
Sample
Control
Seismic
Source
USDC
510
<
10
.4
Small/
Short
depth
slow
little
yes
In
Development
(date
unknown)
yes
SATM
70
80
32.5
100 g/ slow
112m
required
no
Developed &
Operational
no
Gopher
-
-
-
>100g/ slow
1-3m
little
yes
In
Development
(date
unknown)
yes
Our Choice: USDC with Integrated Sensors
Drill Conclusions
Why we chose the Ultrasonic Gopher:
- Low axial force
- Small, light weight, power efficient solution
- Ability to take core sample and in-situ analysis
- Seismic source: mapping of comet
- Integrated Sensor ability:
*accelerometers, fiber optic probes, temperature
sensors and electrodes
- Exciting New Technology with amazing potential
Seismic Sources
• Seismic sources
– Accelerated weight drop system
– Vibrators
– Drill
Seismic Sources
Frquency (Hz)
Accelerated Weight Drop System
Vibrators
7-250
Drill
60-100; 20,000
Best
Worst
Weight (kg)
45-900
20,000
0.3
Power (kW)
15
130-175
.003-.01
Cooling System Requirements
• Must cool to different temperatures
– Be able to cool the sample to temperatures
between 50-200K, depending on temperature upon
acquisition
• Must provide adequate cooling at all
temperatures
• Must operate within mass, power and volume
constraints
Cooling System Trade Study
Cooling System
Ball E100
Ball E200
Sunpower M77
Sunpower M87
radiant cooling
Mass (kg) Cooling Power Power Consumption Radiation
Flight
at 70K (W) at 70K (W)
Hardness (krad) Tested
4.1
2
60
Yes
6.5
3
70
250 Yes
3
3
60
Yes
2.7
4
100
No
8
2
0
No
Best
Worst
In-situ Analysis
•
Scientific Objective
– The instruments chosen need to determine the
elemental, molecular, mineralogical, and isotopic
composition of the cometary surface and subsurface.
– They need to measure properties like nearsurface strength, density, texture, porosity, ice
phases and thermal properties of the comet.
Instrument Description
• COSAC
– (Cometary Sampling and
Composition
experiment)
– Evolved gas analyzer
that detects and identifies
complex organic
molecules from their
elemental and molecular
composition.
Instrument Description
• MUPUS
– (Multi-Purpose Sensors
for Surface and
Subsurface Science)
– Uses sensors on the
Lander’s anchor, a probe
and sensors on the
exterior of the Lander to
measure the density,
thermal and mechanical
properties of the surface.
Instrument Description
• MODULUS Ptolemy
– (Methods Of
Determining and
Understanding Light
elements from
Unequivocal Stable
isotope compositions.)
– Evolved gas analyzer,
which obtains accurate
measurements of isotopic
ratios of light elements.
Instrument Description
• ROMAP
– (Rosetta Lander Magnetometer and Plasma
Monitor)
– Studies the local magnetic field and the
comet/solar-wind interaction.
• Permittivity Probe
– Performs an electrical examination of the cometary
surface layer.
Conclusions
• A sonic drill with encapsulation capability
• The drill will be used as a seismic source
• The Sun power, Inc. M77 cryogenic cooler
attached to an encapsulating drill
• Five instruments for in-situ analysis
•
•
•
•
•
COSAC
ROMAP
MUPUS
MODULUS Ptolemy
Permittivity Probe
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
Orbiter Sensors and Sample Collection
Bautista Fernandez
Edison Estacio
Thomas Jun
Orbiter Sensors & Data Collection
Objective
• Map P2/Encke to acquire images in 3D to help
determine landing site.
• Sense where the space craft is relative to comet
nucleus.
• Sense molecules and other isotopic
compositions of volatile components from
both the cometary coma and nucleus.
Requirements
• Map comet surface for landing
– 1-10 m resolution from safe distance
(est. 30-100 km)
– Determine surface roughness from 1m to 10m
• Monitor distance to comet
– Measure distance within 1 to 10 m accuracy
• Sense molecules and other science information
– IR , UV, and Ion observations
Remote Sensors
• Imaging systems
• Spectrometers
– Thermal Infrared
– Ultraviolet
– Ion Mass
• Heterodyne radio receiver
• Radar altimeter
Imaging System - OSIRIS
Optical, Spectroscopic, and Infrared Remote
Imaging Systems (OSIRIS)
• Used on the Rosetta orbiter
• Consists of two CCD cameras
– Narrow angle camera (NAC)
– Wide angle camera (WAC)
Imaging System - OSIRIS
• NAC is designed to
have high spatial
resolution to look at
comet’s nucleus.
• WAC has a wide field
of view (FOV) and a
high straylight rejection
to look at the dust and
gas above the surface of
the comet’s nucleus.
OSIRIS - Distance Requirement
NAC
WAC
Ground Resolutions (m)
Height (km)
Height (km)
1
50
10
2
100
20
3
150
30
4
200
40
5
250
50
6
300
60
7
350
70
8
400
80
9
450
90
10
500
100
Imaging System - MSI
Multi Spectral Imager (MSI)
• Used on NEAR Shoemaker
• Provides visible and infrared images
• At 12-bits per pixel, an uncompressed image is
1.6 Mbits
Imaging Systems
MSI
OSIRIS
Total mass 9.55 kg
Total Power: 13.04 W
Total mass 23kg.
NAC
WAC
14 microns
14 microns
Pixel Size
-
Array
244 X 537
Focal Length
(mm)
168
700
140
Resolution
95 X 161 mrad
20 µrad/px
100 µrad/px
FOV
2.25° x 2.9°
2.35° x 2.35°
12.1° x 12.1°
Wavelength
400-1100 nm
250-1000 nm
250-1000 nm
2048x2048 pixel x pixel 2048x2048 pixel x pixel
Problem with Requirements
• Resolution of the MSI instrument is too coarse
to satisfy requirements.
• For MSI to obtain a surface resolution of
1-10m, it has to be a distance of 10.5-105.3 m
from the comet.
Preliminary Molecule List
Electromagnetic Spectrum
Wavelengths
Ultraviolet
Infrared
400 nm-10 nm
700 nm-1000 um
Spectrometer - TES
Thermal Emission Spectrometer (TES)
• Used on Mars Global Surveyor
• Measures incoming infrared and visible energy.
• Measures broadband solar reflectance and
thermal emittance
Thermal Emission Spectrometer (TES)
FOV
8.3 mrad per 3x2 array
Spatial Resolution
3 km from MGS orbit
Wavelength Range
6 - 50 microns
Broadband solar reflectance and
thermal emittance
0.3 - 2.7 µm and 4.5 - 100 µm
Mass
14.4 Kg
Average Power
14.5 W
Problem with Requirements
• The wavelength of TES must be modify to
meet Infrared requirement (2.5-50 microns).
Spectrometer - Ultraviolet
Ultraviolet imaging spectrometer (ALICE)
• Used on the Rosetta orbiter
• Obtains far UV-spectra, FUV, multispectral or
monochromatic images of the comet
• Allows analyses of FUV properties of the
nucleus and solid grains
• Is suited to determine abundances of He, Ne,
Ar, Kr, H2O, CO, and CO2
Spectrometer - Ultraviolet
F/3 Primary Mirror
0.1x6 deg2
Bandpass
700-2050Å
Spectra Resolution
5-6Å pt. source
12.5Å extended source at
700Å
9.8Å extended source at
2050Å
Spatial Resolution
0.1x0.6 deg2
Active FOV
0.1x6 deg2
Normal Efficiency Area
0.03-0.53cm2
Mass
2.2 kg
Power
2.9 W
Spectrometer -IMS
Ion Mass Spectrometer (IMS)
• Used on Giotto
• IMS contains two detectors
– High-Energy-Range Spectrometer (HERS)
– High-Intensity Spectrometer (HIS).
Ion Mass Spectrometer (IMS)
Energy Range
20eV to 16keV
Mass Resolution
≈ 20
Mass
9 kg
Average Power
6.3 W
Heterodyne Radio Receiver
Heterodyne radio receiver (HIFI)
• Has continuous coverage in the frequency of
492-1113ghz
• Provides high resolution spectroscopy
(R=104-107) over the frequency interval
480-1170ghz (250-625microns)
• Provides the out-gassing rate of the comet’s
H20 rotational lines
Heterodyne Radio Receiver
Radar Altimeter
TOPEX/Poseidon
• Measures topographical inclinations and
altitudes of surfaces
• Dual frequency (C- and Ku-Band)
• Poseidon can measure the surface height
within 4.3 centimeters.
Poseidon
Frequencies
Mass
Power
13.6GHz and 5.6GHz
23kg
49W
Summary & Conclusions
• Candidate imaging system MSI does not have
sufficient resolution that meets the 1 to 10m
requirement
• Candidate spectrometers (TES) must be modify to
meet Infrared requirement (2.5-50 microns)
• Heterodyne Radio Receiver does not meet
mass and power requirements
• Imaging System - OSIRIS
• Spectrometers - Ion Mass, TES, ALICE
• Altimeter - TOPEX/Poseidon
Outstanding Issues
• A light Heterodyne Radio Receiver that
consumes minimal amounts of power needs to
be researched
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
Smart Dust
David Ibañez
Vincent Lin
Smart Dust
Objective:
• To investigate applications of miniature
sensors in a wireless network (Smart Dust)
useful for our CSR mission.
• Plan out possible ways to complete proposed
applications.
Smart Dust
Outline of Presentation
•
•
•
•
•
•
•
•
•
Introduction on Micro Motes
Seismic Application
Accelerometer Application
Landing Site Application
Wireless Mote Trade Study
Power, Mass, and Volume
Deployment
Anchoring
Conclusion
Introduction to Smart Dust
Introduction:
• Currently under research at Berkeley research
labs
• Millimeter-scale wireless sensor devices
capable of collecting various chemical and
physical stimuli.
• Micro Motes built from off-the-shelf
components
RF Micro Mote
Seismic Application Illustration
Seismic Application
• Develop an 2D image Comet’s internal structure
• Seismic Tomography
– Seismic Source for sending waves into the comet
– Collect data with Seismic sensors
• Requirements
– Good sensor coverage on comet surface.
– Appropriate seismic source and sensors(low frequency <500Hz).
Seismic Tomography
• Earth's structure at 200
km depth below
Southeast Asia
• Blue color area
represent high seismic
velocity
• Red color area represent
low seismic velocity
Seismic Sensor
Geophones
Freq. (Hz)
Sample Rate (ms)
Mass (g)
Volume (cm3)
SM-15
14
1
74
16.214
SM-24
10
2
74
16.214
Seismic Formation
Accelerometer Application Illustration
Accelerometers
Trade Study of Accelerometers
Accelerometers
Power
Range
Dimensions
Axis
Consumption
Bandwidth
Analog Devices
ADXL202
2.7~5.25 V
±2g
5x5x2mm
dual
0.5uA
10hz
Analog Devices
ADXL311
3V
±2g
5x5x2mm
dual
0.4mA
10hz
Analog Devices
ADXL210E
3~5v
±10g
5x5x2mm
dual
0.6mA
10hz
Reiker B3
3~6V
±3g
25 dia.x11mm
dual
~1mA
0~160hz
Fujikura 3 axis
5V
±2g
5x5x2.6mm
tri
<50mA
0~150hz
Sislicon VSG
CRS03-11
5V
±10g
27x13x27mm
tri
<50mA
10hz
Landing Application Illustration
Landing Guidance Application
• Deploy Optical Motes to potential landing site
• Attempt optical communication between
motes
– If level land site, optical communication are
expected to be functional
– If there are obstacles, optical communication fails
because of direct-line-of-sight requirement
• Optical communication may fail due to debris.
(Need to use Radar Altimeter)
Mote Comparison
Trade Study of COTs Dust
Mote
Type
Transmission
Range
Data
Transfer
Rate
Transmission
Properties
Dimensions
Mass (g)
RF
3 - 200 m
RFM
Transceivers
1 - 20
kbps
Laser
~10's km
CCR
IrDA
Sleep/Wake
Current@3volt
300 - 916.5
MHz
75x25x12.5mm
~150
1 + 0.75uA/
7 + 8mA
4 bps
650 nm
25x25x51mm
~160
1uA/25mA
~150 m
30 bps
Passive
Laser
18x18x2.5mm
~160
50uA/10mA
~60 cm
115 kbps
800-300 nm
50x11x11mm
N/A
RF Mote battery life (3V Lithium ion battery): 66 hours
continuous operation, 1.5 year at 1% duty cycle.
Transceiver
• RF Monolithics TR1000
– Greatest transmission
range (3~200m)
– Lowest current
consumption
– Data transfer rate of
20Kbps
• Requires 7.79 cm
vertical antenna
Power, Mass, and Volume
Subject
RF Mote
Power (mW) Mass (g) Volume (cm3)
2160
7200
1180.8
Geophone
SM15
Seismic App.
0
5920
1297.12
2160
13120
2477.92
Accelerometer
2217.6
7200
1183.2
Landing
N/A
N/A
N/A
Deployment
Canister
32 motes
22.6 cm dia. x 7.62cm
16 motes
11.3 cm dia. x 7.62cm
Anchors
• Multi-leg
• Best stability
• Mounted on sides
– Mote
– Geophone
• Compatibility with
mode of deployment
Conclusion
• Three applications for the Smart Dust team
– Seismic, Accelerometer, and Landing Guidance
• Cannot do Landing Guidance
• Depending on Power, Mass, and Volume
constraints set for the wireless mote
applications, we may not be able to do both the
Seismic and Accelerometer applications.
• We need to determine the trajectory of the
Canister deployment methods.
Overview
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and
Communications
• Spacecraft Computer System and Networking
• Space System Integration
• Summary and Conclusions
Data Collection, Compression, and
Communications
Rinesh Patel
Parisa Toorani
Data Collection and Communications
Objectives
• Design communication links to transmit data
– Between the sensors, the probes, the
Lander and the orbiter
– Between earth and the orbiter and between
earth and the Lander
• Theoretical frequency and data rate analysis
Link Considerations
• High gain antenna for space-to-earth link for
the Lander and the orbiter
• Low gain antenna for communication between
the probes, the orbiter, and the Lander
• Ka-band links between the orbiter and earth
and between the Lander and earth
• UHF links between the space modules
• NASA Deep Space Network (DSN)
Communication Phases
Phases of mission:
• Cruise – high gain antenna communication
(4 hour access time/day)
• Reach Encke – orbiter maps the comet and
transmits the images via a high gain antenna
(10 hours continuous access time/day)
Communication Phases (cont.)
• Landing
- Lander communicates with the probes and
orbiter via low gain antennas
- orbiter and Lander communicate with earth
(4 hour access time/day for each)
• Return trip – Lander and orbiter continue
transmitting data (3 hour access time/day for
orbiter, 5 hour/day for Lander)
Link Budget
Signal to Noise Ratio:
Eb/No = (P • Ll • Gt • Ls • La • Gr)/(k • Ts • R)
•
•
•
•
•
•
•
•
•
P – Power (W) = 12 W
La – Atmospheric Loss (dB) ≈ 0.9
Ls – Space Loss (dB) = (λ/(4π(4.5x1011 m ))2
Ll – Link Loss (dB) ≈ 0.9
Gt – Transmitter Gain (dB) = 4(π•r)2/λ2
Gr – Receiver Gain (dB) = 4(π•17.5 m)2/λ2
k – Boltzmann’s constant = 1.38x10-23 J/K
Ts – System Temperature = 100 K
R – Data Rate (bps)
Data Rate Estimation
Eb/No = 2.7, Receiver Ant - 34 m, 25 GHz, BPSK Reed-Solomon w/ Viterbi
coding
Data Transmission Time
•
•
•
•
•
Image file 2048 x 2048 pixels (24-bit) ~ 12.3 MB
Time of propagation ~ 30 min
Data Rate ~ 23 kbps (60 cm diameter Antenna)
Send one image (uncompressed) ~ 2 hr
4:1 compressed image (3 MB) ~ 48 min
Image Resolution
Wide Angle Camera
Narrow Angle Camera
Distance
(km)
FOV (m)
Resolution
(m/pixel)
Distance
(m)
FOV (m)
Resolution
(m/pixel)
10 km
2119.73
1.03
10 km
410.2
.2
20 km
4239.46
2.07
20 km
820.42
.4
30 km
6359.2
3.11
30 km
1230.6
.6
40 km
8478.9
4.14
40 km
1640.8
.8
50 km
10598.7
5.17
50 km
2051.05
1
# Frames vs. Data Size
Narrow Angle Camera
Wide Angle Camera
FOV
# of
Data
FOV
# of
(m) Frames
Size
(m)
Frames
410.2
75
225 MB 2119.73
3
820.42
19
57 MB 4239.46
1
1230.6
8
24 MB 6359.2
1
1640.8
5
15 MB 8478.9
1
2051.1
3
12 MB 10599
1
Data
Size
9 MB
3 MB
3 MB
3 MB
3 MB
Total Estimated Data
• Each camera mapping Encke 10 times
40 Uncompressed Images (2048 x 2048) ~
492 MB
• Housekeeping ~ 49.2 MB
• Sensors ~ 49.2 MB
• Total Data ~ 591 MB
• Sufficient storage space
X- and Ka-Band Frequencies
Frequency Ranges*
Frequency
Band
Transmit f
Receive f
DSN Operating
Frequencies
DSN
Uplink f
DSN
Downlink f
x (MHz)
7145 –
7190
8400 –
8450
7900 –
8400
7250 –
7750
ka (GHz)
34.20 –
34.70
31.80 –
32.30
27.50 –
31.00
17.70 –
19.70
* Regulated by International Telecommunications Union (ITU) and the
World Administrative Radio Conference (WARC)
Atmospheric Attenuation
This plot shows the attenuation of electromagnetic waves as
they enter from different heights above the atmosphere.
NASA Deep Space Network (DSN)
• Three stations each
separated by 120°
longitudinally
- California
- Australia
- Spain
• 26 m, 34 m, and 70 m
antennas
• Estimated 10 hour time
slot per day
Picture obtained from JPL
Data Compression
Lossless Compression
• Original data can be
recovered almost
entirely
• Does not work if
original data is analog
• Compression ratio
varies from 2:1 to 4:1
Lossy compression
• Original pixel intensities
can not be recovered
• Can reconstruct most of
the original image
• Compression ratio
varies from 4:1 to 40:1
Data Compression Trade Studies
+ indicates rating
Information obtained from EV-3M project
Antenna Trade Studies
Manufacturer
Performance
COM DEV
Ltd.
23 GHz (180 MHz bandwidth)
36 dBi gain
COM DEV
Ltd.
19 GHz (190 MHz bandwidth)
27 dBi gain
Power
Size
Mass
5.5 kg
0.254 m 5.44 kg
diameter,
0.305 m
high
Transponder Trade Studies
Manufacturer
Performance
Motorola
Ka-band (31.8-32.3 GHz)
downlink
X-band Uplink (7.1457.235 GHz)
Output power: (x) – 12
dBm, (ka) – 4 dBm
12.9 W
Frequency Range 17.721.2 GHz
Output Power: 26 dBm
max
< 11 W
Alcatel
Power
Size
Mass
3 kg
201 x 66 x < 1.25 kg
152 mm
Amplifier Trade Studies
Manufacturer
Motorola
SSSD
Performance
3.4 W output
23 GHz
Power
Size
19 W
10.2 x
22.9 x
2.54 cm
AIL Systems
Inc.
13.8 dBm output per phase
27.5 GHz
18.3 dB nominal gain
13.5 W
3.76 x
2.49 x
0.305 cm
Alcatel
2 W output
17.7-20.2 GHz (500 MHz
BW)
30 dB nominal gain
12 W
60 x 60 x
15 mm
(for 2)
Mass
75 g
Conclusion
The following are being considered:
•
•
•
Ka-band frequency 20–25 GHz link
UHF links
JPEG 2000 data compression
Outstanding Issues
• Find equipment to closely match our
expectations
• Decide how much data to send/store
• Time scheduling between orbiter-to-earth and
Lander-to-earth communication links
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
Spacecraft Computer Sys.
Turtle Kalus
Tarik Elsorady
Jeffry Gosal
Why do we need a computer?
•
•
•
•
Communications
Instrument and Sensor Control
Signal and Data Processing
Data Compression
• Navigation and Attitude Control
• Power and Resource Management
Objectives
• Design a Computer Control System which can
handle our mission requirements
• We want to optimize:
– Capability
– Flexibility
– Reliability
Computing Requirements
•
•
•
•
•
Speed / MIPS
Power Consumption
Networkability
Memory Capacity
Built-in Error
Correction
• Radiation Hardened
•
•
•
•
•
Latch up Immunity
Reliability
Cost
OS Compatibility
Smallest Possible
Package
• Redundancy
Central Processors
• Processors
–
–
–
–
–
Aitech S220
Aitech S320
Honeywell Space Computer
BAE Systems RAD750
Seakr PPC 603e
Central Processor Comparison
• Computer Systems
Aitech S220
Mass
N/A
MIPS
N/A
Power (W) nom./max.
15 / 17.5
Radiation Hardness (kRad) 20
Redundancy
single
Reliability / years
.995 / 25
Boot/User ROM
2MB / 64MB
Speed (MHz)
133
SRAM
256MB
Temperature Range (°C)
-40 to +71
Better
Best
Aitech S320
N/A
N/A
10 / 12.5
20
single
.995 / 25
2MB / 64MB
133
256MB
-40 to +71
BAE RAD750
549 g
21
<10.2
> 100
single
.975 / ~44.5
256kB / 1MB
133
128MB
-55 to +80
Honeywell HSC
4545 g
2.2
<5
> 100
dual
.995 / ~12
N/A
12
512kB
N/A
Seakr PPC603e
< 1000 g
22
<~10
> 100
single
N/A
512kB / 2MB
133
128MB
N/A
Seakr Engineering
PPC 603e Processor Card
Radiation Characteristics
• Radiation and Temperature Tolerant
• Reliable, maintenance-free operation in space
• Built using fully controlled and documented
processes, approved for use in Deep Space
• All components are immune to latch up
• Low Single Event Upset (SEU) rate
Computing Specs
•
•
•
•
•
•
•
PowerPC Compatible RISC µ-processor
Standard VMEbus Interface
222 MIPS @ 133 MHz
Power Consumption < ~10 Watts
128MB DRAM and 2MB EEPROM
Built-in 10Base-T Ethernet Interface
Can be replaced with a PPC750
Memory System
• Magnetic based (ordinary hard drive)
– Highly susceptible to corruption in space
– Bad choice
• Solid State
– Solid state memory better withstands
the radiation environment of space.
Memory
• Memory Systems
–
–
–
–
Aitech S290 16GB High Density Flash Memory
BAE Systems 1GB Solid State Recorder
Chrislin CI-VME64 – 4GB
Seakr 24GB Solid State Recorder
Memory Comparison
• Memory Systems
Aitech 290
Power (W) nom/max
4/5
Radiation Hardness (kRad) N/A
Reliability / years
N/A
Temperature Range (°C)
-40 to +71
Memory Capacity
128MB to 16GB
Better
Best
BAE SSR
N/A
> 60
NA / ~44
-30 to +65
1GB
Chrislin CI-VME64
10 / 12.5
N/A
N/A
0 to +50
32MB to 1GB
Seakr 24-GB VME
N/A
N/A
N/A
-50 to +70
24GB
Seakr Engineering
24GB VME Memory Card
Seakr Engineering Memory Board
• Seakr Engineering Solid State Memory Card
–
–
–
–
–
–
–
–
Standard 6U VME Form Factor
Total Storage: 24GB
Power Consumption: 4 to 6W
Operating Temperature Range: -50°C to +70°C
30 MB/sec transfer rate
Non-Volatile Flash Ram
Proprietary Seakr EDAC software available
> 70,000 hours between failures *
Seakr Engineering
Command and Data Handling System
Seakr Engineering C&DH Sys
• Command and Data Handling System
– Standard VME backplane
•
•
•
•
Widely used form factor
Many COTS applications
Saves money in development
Expands selection for other peripherals
– Readily Available from Seakr
• Assures compatibility
Seakr Engineering
Hardware Image Compressor
Seakr Engineering Image Compressor
• JPEG2000 Hardware Image Compressor
and Data Buffer
–
–
–
–
–
High-speed real time compression
Compression ratios greater than 80:1
Flexible quality, resolution and color
Solid state memory storage
Frees up main processor(s)
Seakr Engineering
Re-Configurable Computer (FPGA Board)
Seakr Re-Configurable Computer
• FPGA Board
–
–
–
–
–
Four Reconfigurable Co-Processors
1 GB High-Speed SDRAM Local Memory
Processor Controlled Reconfiguration
SEU Fault Tolerant
Automatic SEU Detection and Scrubbing
Equipment Overview
• Orbiter
–
–
–
–
2 - PPC 603e boards
1 - 24GB SSR
1 – FPGA board
1 – Image Compressor
Board
– 1 – C&DH System
• Lander
–
–
–
–
2 - PPC 603e boards
1 - 24GB SSR
1 – FPGA board
1 – C&DH System
If Image Compression is
needed aboard the
Lander, use FPGAs
Computing Redundancy
• Orbiter
– Hardware
• Cold Standby Spare
during Cruise Phase
• Hot Standby Spare and
using Seakr’s FPGA
board for processor
intensive calculations
during Intercept and
Landing Phase
– Software
• Use Seakr’s built-in
EDAC system
• Lander
– Hardware
• Complete Shutdown
during Cruise Phase
• Same configuration as
Orbiter for Intercept,
Landing Phase and Return
Phase.
– Software
• Use Seakr’s built-in
EDAC system
Memory and Data Redundancy
• Orbiter and Lander
– Two duplicate sets of data on both the Orbiter and
the Lander.
– Encode data using Hamming code.
Computer Systems Block Diagram
24GB VME
Memory
FPGA
Board
PowerPC
603e card
1
PowerPC
603e card
2
Image
Compressor
& Data Buffer
Only in Orbiter Comp.
Communications
link
Embedded
Systems
Mission Data-Link Phases
Comet
Orbiter
Lander
Separation
Landing
Phase
Intercept and
Mapping
Phase
Cruise Phase
Sample
Collection
Phase
Lander Return
Phase
Probe Data
Collection
Phase
Real Time Operating System
•
•
•
•
Small Memory Footprint
Short Interrupt Latency
Predictable Behavior
Designed for Reliability and Critical
Applications
Real Time Operating Systems
•
•
•
•
•
Integrity
LynxOS 4.0
QNX 6.2
VxWorks 5.x
VxWorks AE
RTOS Comparison
VxWorks AE
Support for PowerPC family
Yes
Support for Protection Domains
Yes
Number of Priority Levels
256
Maximum number of Tasks
Limt'd by mem.
Kernel ROM (min, max)
N/A
Kernel RAM (min, max)
N/A
Minimum RAM per process
N/A
Minimum RAM per thread
N/A
Minimum RAM per queue
N/A
Typical thread switch latency
N/A
Guaranteed max. inerrupt latency N/A
System Clock Resolution
N/A
Priority Inversion Avoidance Mech. Yes
Multiprocess Support
Yes
Multiprocessor Support
Yes
VxWorks 5.5
Yes
No
256
Limt'd by mem.
N/A
N/A
N/A
N/A
N/A
N/A
N/A
N/A
Yes
Yes
Yes
Inegrity
Yes
No
255
N/A
70K
20K
1 Page
1 Page
128 bytes
50-100clks
N/A
N/A
Yes
Yes
Yes
LynxOS 4.0
Yes
No
256
Limt'd by mem.
280K, 4M
500K, 4G
1073 bytes
1073 bytes
80 bytes
4us to 19us
14us
20us
Yes
Yes
Yes
QNX 6.2
Yes
Yes
64
Finite
64K, 64K
N/A
N/A
N/A
N/A
depends
depends
depends
Yes
Yes
Yes
VxWorks AE
• Derived from the widely adopted VxWorks
5.X RTOS
• PowerPC support
• Many network protocols supported
• Implements “protection domain” technology
which effectively separates resources and
helps prevent conflicts
Protection Implemented in VxWorks AE
• Protection Domain
Technology
– Single, flat physical
address space is
extended to multiple
partitions
– Enables developers
to create logical
“containers” to
isolate and protect
applications
Selection
• Computer Systems
–
–
–
–
4 Seakr PPC 603e
2 Seakr SSR - 24GB
2 Seakr C&DH System
1 Seakr JPEG2000 Image Compressor
(only aboard Orbiter)
– 2 Seakr RCC – FPGA Board
– VxWorks AE
Outstanding Issues
• Networking
• Cost Analysis
• Other teams computing needs
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
Space System Integration
Paul Mellentine
Richard Harris
Space System Design Integration
• Objective
• UCSC integration with University of Michigan
• U of Michigan PDR Results
– Power System
– Propulsion
• Launch Vehicle
• Spacecraft
• Mass Itemization
Objective
– To integrate the UCSC payload designs with the
University of Michigan spacecraft design
– To keep the UCSC design within power, mass, and
volume constraints
UCSC and University of Michigan Joint Effort
UCSC
U of M
• Mission Science &
Architecture
• Structures & Materials
• Nucleus Sample
Return & Analysis
• Power & Thermal
• Remote Sensing &
Analysis
Complete
System
Design
• Trajectory & Propulsion
• Smart Dust
• Guidance, Navigation,
Command, Control, &
Communications
• Communications
• Test & Quality Control
• Payload Computers
• Budget & Scheduling
Joint Tasks:
Mission Science
Communications
Teaming with U of Michigan
• Prelimary Study
Teaming with U of Michigan
• Aerospace Engineering 483
– Cometary Science and Sampling Endeavor (CSSE)
• Preliminary Design Review
Differences with Michigan
• Michigan Mission Architecture
– Comet Wild 2
5.2 AU
– Orbiter & Lander Return
• UCSC Mission Architecture
–
–
–
–
Comet Encke
Lander Return
Smart Dust
Dust Probes
2.2 AU
Orbiter Payload Power
Mission Phases
Orbiter
initial trip mapping landing
drilling &
take-off return trip
analysis
Command and Data Handling
30
30
30
30
30
30
Telecommunications
4*
4*
4*
4*
4*
4*
Cameras
7*
17*
17*
7*
7*
7*
Spectrometers
10*
10*
10*
17*
10*
10*
Radar Altimeter
10*
49
49
49
49
TOTALS (W):
51
110
110
107
100
51
* Estimate from U of M CORSAIR
Power requirement ~ 110 W
Lander Payload Power
Mission Phases
Lander
initial trip mapping landing
drilling &
take-off return trip
analysis
Command and Data Handling
30
30
30
30
30
30
Telecommunications
4*
4*
4*
4*
4*
4*
Narrow Angle Camera
7*
15*
15*
7*
7*
Instruments
10*
10*
10*
17
10*
10*
Drilling Device
10
Sample Containment
10
10
10
100
100
100
TOTALS (W):
61
69
69
168
151
149
* Estimate from U of M CORSAIR
Power requirement ~ 170 W
Propulsion
NASA 457-M Hall Thruster
Advanced Concept Ion Thruster
•
•
•
•
•
•
•
Use: Earth escape/capture, comet
rendezvous
Operating conditions:
- 58.5 kW nominal power
- Max thrust = 2.5 N
- Discharge Isp = 2995 s
Will provide a total ∆V of 4.55 km/s
Total burn time: 0.54 yrs = 198 days
Required Xe Propellant: 1454 kg
Total system dry mass: 344 kg
•
•
•
•
•
•
•
•
•
Use: Deep space cruise between Earth
and Wild-2
System will employ 4 thrusters with 3
thrusters in operation at all times
Operating conditions:
- 3*20 = 60 kW nominal power
- Max thrust = 3*0.400=1.200 N
- Discharge Isp = 7500 s
Will provide a total ∆V of 14.89 km/s
Total burn time: 2.44 yrs = 892 days
Required Xe Propellant: 1258 kg
Total system dry mass: 556 kg
Source: University of Michigan Cometary Science and Sampling Endeavor PDR
Power Generation Systems
Advantages
Solar
Radioisotope
• Unlimited Power
Source
• Low fuel
consumption
• Not reliant on the
sun
• Unlimited life
• High output to
• Long discharge
fuel ration
life
• Produces radiation
• Heat radiator
required
• Produces
radiation
• Requires
storage of water
and gases
25 – 200
5 – 20
2 – 40
275
0.2 – 300
0.2 – 10
5 – 300
0.2 – 50
• Poor at large
distances from the
sun
Disadvantages
• Cells can be
damaged by comet
dust tail
Specific Power*
(W/kg)
Power Range*
(kW)
Nuclear
Fuel Cell
Source:Wertz and Larson, ed. Space Mission Analysis and Design. Third Edition
Power Systems
Orbiter Power
• Radioisotope thermoelectric generators (RTGs)
• Power output for two: 20 kW
• Specific Power: 65.09 W/kg
• T/E conversion: 7% efficiency
• Power system mass: 307.22 kg
Lander Power
• Heat-pipe cooled nuclear reactor
• Power output: 61 kW
• Specific Power: 17 W/kg
• T/E conversion: Brayton cycle (27% efficiency)
• Power system mass: 3,588 kg (10% margin)
Launch Vehicle
• Phase One
– Liftoff from Earth
– Escape from Earth’s sphere
of Influence
• What Launch Vehicles will
work for our mission?
– Mass
– Volume
Spacecraft Mass Itemization
Lander
Command and Data Handling
Telecommunications
High Gain Antenna
Low Gain Antenna
Transponders
Amplifier
Wide angle camera *CORSAIR
Laser Range Finder *CORSAIR
Payload Instruments
COSAC
ROMAP
Modulus Ptolemy
MUPUS
Permittivity Probe
Drilling Device (USDC)
Crytogenic Cooler
Samples (3)
Guidance, Navigation, & Control *
Electronic Propulsion System *
Thermal Control **
Structure **
Intial Estimate
30% Design Margin
Total
Mass (kg)
12.000
5.500
5.440
1.250
0.075
13.300
2.160
4.850
0.700
4.500
4.500
0.500
0.400
3.000
0.300
30.108
3158.500
231.93
1159.67
4638.690
1391.607
6030.297
Orbiter
Command and Data Handling
Telecommunications
High Gain Antenna
Low Gain Antenna
Transponders
Amplifier
Cameras (OSIRIS)
Spectrometers
Ion Mass
Thermal Ir & Visual
UV
Radar Altimeter
Probes
Smartdust
Guidance, Navigation, & Control *
Electronic Propulsion System *
Thermal Control **
Structure **
Intial Estimate
30% Design Margin
Mass (kg)
12.000
5.500
5.440
1.250
0.075
23.000
2.200
14.400
2.200
23.000
150.000
11.200
25.583
453.500
36.467
182.337
729.348
218.804
Total
948.152
Orbiter + Lander
6978.45
307.940
3588.000
RTG Power System
Nuclear Power System (10% margin)*
Spacecraft Total
Delta IV-H Payload
10874.389
10889.000
Mission Launch Vehicles
Mission
Total Mass
Launch Vehicle
CSR
10875 kg
Delta IV-H
Cassini
5712 kg
Titan IV
Rosetta
2900 kg
Ariane V
Deep Impact
1020 kg
Delta II
Giotto
960 kg
Ariane I
Stardust
385 kg
Delta II
Gravitational Assist
Rosetta Mission Example
Integration Team Conclusions
• Power Requirements
– Orbiter Payload
– Orbiter Ion Engine
0.110 kW
20 kW
– Lander Payload
– Lander Hall Thruster
0.170 kW
60 kW
• Power Generation Systems
– Orbiter
• Radioisotope thermoelectric generators
20 kW
– Lander
• Heat-pipe cooled nuclear reactor
61 kW
Integration Team Conclusions
• Mass
– Total Spacecraft
10875 kg (30% Margin)
• Launch Vehicle
– Delta IV-H
• 10899 kg Payload with
Earth Escape
• 4 meter Fairing
Outstanding Issues
• Total Mass
– Too large?
– Explore Alternate Deep Space Travel Methods
• Gravitational Assist
Overview
•
•
•
•
•
•
•
•
•
Introduction
Mission Science and Architecture
Lander Sensors and Sample Collection
Orbiter Sensors and Sample Collection
Smart Dust Applications
Data Collection, Compression, and Communications
Spacecraft Computer System and Networking
Space System Integration
Summary and Conclusions
CSR Objective
• To provide humanity with a better explanation
and understanding of the source of organic and
non-organic molecules and their relation to
Comets and life on Earth.
• To give insight on the evolution of comets
from their origins to their destiny, and to
understand the source and evolution of the
solar system, planets, asteroids, and space
debris.
Scientific Objectives of CSR
• To study:
– Internal Structure
– Ion and Dust tail
– Organic Parent
Molecules
– Rotation
– Activity while
Approaching the Sun
• To Return:
– Pristine Samples of a
Comets Nucleus
– Pictorial Atlas of the
Surface of a comet
– Sensor Data
•
•
•
•
•
Rotational
Spectrometer
Permittivity
Magnetometer
Dust Analyzer
Mission Components
Orbiter
Coma
probe
Lander
Tail
probe
Launch
• Delta IV-H
– Spacecraft mass
10875 (30% margin)
http://www.boeing.com/companyoffices/gallery/images/space/delta_iv/d4_1st_flight_28.htm
Cruise
• Hall Thruster
– 60 kW power
– 1 on Lander
• Use: Earth escape/return
• Ion Engines
– 20 kW power per engine
– 3 on Lander
• Use: Deep Space Cruise
– 1 on orbiter
• Use: Following comet
• Approach Comet
Power
• Orbiter
– 2 Radioisotope
Thermoelectric
Generators
• 20 kW
• Lander
– Heat-pipe cooled
nuclear reactor
• 61 kW
Communications
• Data Rate ~ 10 kbps
• Frequency
– ka band (~ 20–34 GHz)
– Antenna
• Com Dev Ltd. 19 GHz 0.254m
• JPEG 2000 Data Compression
• Alcatel Transponder 26 dBm
• Alcatel Amplifier 30 dB gain
DSN Access Time Estimates
Phase
Use
(hours/day)
Cruise
2
Orbit
10
Landing
7
Return
5
Computer System
• Seakr PPC603e CPU
– 133 MHz, 128 MB SRAM
• Seakr 24 GB VME Solid State Memory Card
• Redundancy
– 4 Processors: 2 Orbiter, 2 Lander
– Two 24GB Solid State Memory: 1 Lander, 1 Orbiter
• Mirrored data storage until Lander returns to Earth
• VxWorks AE
Comet Rendezvous
• Follow Comet at Lagrange Point ~50 km
• Orbiter
– Surface and Nucleus imaging and mapping
• OSIRIS 1m resolution
• Determine landing
– Molecular Composition Analysis
Spectrometers
• TES Thermal Emission
• ALICE Ultraviolet imaging
• Ion Mass
– TOPEX Radar Altimeter
• Topographical map
• Orbiter altitude
Landing
• Deploy Lander
• Drill Nucleus Sample
– Drill at least 1ft deep for
pristine sample using USDC
– Three 100g samples
– Sunpower M77 Cooler
• In situ Measurements
–
–
–
–
–
COSAC elemental, molecular gas analyzer
ROMAP magnetometer and plasma monitor
Modulus Ptolemy isotopic ratio
MUPUS density, thermal properties
Permittivity Probe
Smart Dust
• Small COTS Wireless RF Network Sensors
• Analysis
– Seismic Tomography using SM-15 Geophone
• 48 Motes for half hemisphere coverage &
communication to Lander
– Rotational using ADXL311 Accelerometer
• Canister Deployment
Lander Earth Return
• Return to Earth
– Ion engines
– LEO orbit
• Hall thruster
– Space Shuttle pickup
Probes
• Orbiter deploy probes
–
–
–
–
–
Surface and Nucleus imaging
IR and UV Spectra
Ion mass measurements
Surface mapping
Magnetic field
• Orbiter relays probe data
Why Comet Sample Return
• “Fossils” in revealing primitive composition of
Solar System
– Comet Composition: elemental, isotopic, organic, and
mineralogical measurements
– Understand formation and evolution of Solar System
• Chemistry in extreme physiochemical conditions
Outstanding Issues
•
•
•
•
Lightweight Heterodyne sensor
Computer Networking
Smart Dust Canister Deployment Dynamics
Total Spacecraft Mass
– Gravitational Assist
• Michigan Integration
– Mission Architecture Differences
Final Comments
Comets are sources of valuable scientific data for
clues to understanding the origin of solar system
and origin of life on Earth.
Acknowledgements
Alec Gallimore
Greg Laughlin
Douglas Lin
Claire Max
Graham Smith
Martin Yabroff
Ken Lande