Download Recurring Cost ROM

Advantages of Very
Small Spacecraft
15 May, 2007
Pete Klupar
[email protected]
Definitions
Development
Mass
Large
Cost
Time
2000kg+ 1,000M+ 10yrs+
Small
750kg
100M
2-3yrs
Mini
250kg
75M
2yrs
Micro
100kg
50M
1.5yrs
Nano
1-10kg
5M
~1 yr
Pico
100gm
> 500k
months
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Without
Proposed Permission
By Surrey Satellite Technology Limited
ARC Small Spacecraft Division
• Develop Sustainable Cost Effective Space Missions To
Enable Access To Space
• Common, Reusable Architectures
– Emphasis On Payloads And Science
• Provide Space Access that is Reliable, Frequent and Low
Cost
– Small Space Systems
– Secondary Payloads
• Reduce Overall Mission Costs
– Goal: Maintain Or Increase Scientific And Exploration Return
While Reducing Life Cycle Costs
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Small Spacecraft Projects
• GeneSat and GeneBox (Flown)
• Lunar Science Orbiter
(LSO -Proposed)
• Common Bus
(Lunar Lander Concept Shown)
• Lunar Crater Observation
Sensing Satellite
(LCROSS - in development)
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Background – International Activities
Country/Entity
Small Satellite Programs
United Kingdom
SSTL ~ 40 missions <$100M, 5-500Kg; DERA/QINETIQ (STRV)
ESA
Smart-1, PROBA-1, PROBA-2……PROBA-N
France
CNES - Myriade <150kg S/C, <70kg P/L, 6 launched since 2004, 10 in
development
Japan
JAXSA – Index (72 Kg, 2005 launch <$10M)
Sweden
Swedish Space Corp – 6 Small/Microsats in orbit, 3+ in development
(Viking, Freja, Astrid 1,2 Odin, Prisma, Svea etc)
Germany
DLR, TuB (TUBSAT-A, -B, -N/N1,-DLR, -MAROC,- LAPAN)
Denmark
DTU, Terma – Oerstad, Romer
Israel
Rafael, IAI – EROS-A, EROS-B (Imaging Microsatellites)
Canada
Dynacon/UTIAS – MOST, NESS, Brite, MDA – Rapid Eye
India
ISRO – HAMSat (45 kg microsatellite)
Others
China, South Africa, Turkey, Chile, Nigeria, Korea, Taiwan, Australia, Eqypt,
Indonesia, Russia, Malaysia, Belgium
International Efforts Include >1000 Small Satellites
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Emerging Small ELVs Offer Cost Effective Performance
Launch Vehicle
Pegasus
200 km, 38 °
LEO Mass (kg)
Estimated
425
GTO Mass (kg)
Estimated
N/A
TLI Mass (kg)
Estimated
N/A
NEO Mass (kg)
Estimated
Fairing Diameter
(m)
Price ROM ($M)
N/A
1.3
$35
Taurus 3110/3113
1530
627
427
?
1.6
$50
Taurus 3210
1291
107
367
N/A
2.3
$50
Minotaur 1
565
N/A
1.3
$25
Minotaur 4/5
Falcon 1
N/A
N/A
1700
692
464
425
2.3
$25 - $31
570
107
82
N/A
1.5
$10 - $12
Unique
opportunity
for increased
mass at
substantially
lower cost
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Minotaur V - Star 37GV
Composite Clamshell Fairing
• Flight Proven 92” Taurus Design
Stage 5 Assembly
• Star 37GV Solid Rocket Motor (New for
M-V)
– Thrust Vector Controlled
•
OSP-Standard Avionics
•
•
Cold Gas Attitude Control System (ACS)
Composite Structure
– Only Subset Required to Fly Stage 5
Guidance Control Assembly (GCA)/Stage
4
• GCA Design Shared with Minotaur III & IV
• OSP-Standard Flight Proven Avionics
– Split Between S4 and S5
• Performance:
• Cold Gas ACS
– 496 Kg to TLI
• Stage 4 Star 48V SRM (New for M-V)
– Thrust Vector Control
– Qualified via Static Fire
•
Total Launch Cost (ROM):
– ~$36M (First Mission)
GFE Peacekeeper Stages
• Includes S-37GV Qual
• Stage 3 - SR120
– ~$26M (Recurring)
• Stage 2 - SR119
• Stage
1 -SR118
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Significant Excess Performance
• Launch Vehicles
Provide Hundreds Of
Kilograms Of Excess
Performance Yearly
• Effective Space Exploration
Requires Continued
Development And
Demonstration
• This Requires Routine, Low
Cost Access To Space
• Opportunities For 6 To 12
Secondary Payloads Per
Year
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Notional Costs and Schedule
MITEC
$10M
NFIRE
Optimized Design
Based on existing
Instrument
SPARE with
New 7.5cm
Telescope
MSO
SPARE
with New
30cm
Telescope
Budgetary
Cost ($)
SIS/SWIMS SPARE
MSTI-3
SPARE
MISTEC with
16cm
Telescope &
New FPA
$5M
EPAM(as is)
SWEPAM with
upgraded sensor
6
12
Payload Delivery or Availability Schedule
(months)
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18
Recurring Cost ROM
• Overall Recurring Goal For 5th Unit Is $2.0 M
• Major Recurring Cost Drivers
–
–
–
–
–
–
Communication Equipment $800K To$1m
Radiation Hard Computer: $400K
Star Tracker Equipment: $200K
Propulsion System: $150K
Assembly And Testing: $150K
Telescope System $100K
• COTS Components Vs Space Qualified
Components
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Small Concept
Star Tracker
Patch
Antennas
Diplexer
Transmitter
Receiver
Amplifier
Avionics
Additional payload space
as available
Battery
North side panel for
externally mounted
payloads
Radar
Altimeter
DSMAC
Payload(s) located
internally
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Small Lander Payloads
Stereo imaging system
Surface images for analysis
Mast for stereo imager
Provide elevation for imaging
Mass Power
Duty Cycle
(kg)
(W)
0.8
6.0 360° images
3.5
9.5 1 deployment
Drill, deployment mech
samples from depths of 2 m
20.0
30.0
Determine volatile compounds and isotopic composition
19.0
75.0
Lander Payload Element
Gas Chromatograph Mass
Spectrometer
Sample processing system for
GCMS
Beacon
Objective
2-4 hrs at station
2-hour analysis measurement
For each GCMS sample
Process core or scoop material for analysis.
Navigation reference
1.0
5.0
Magnets
Magnetic susceptibility of regolith particles
0.5
0.0
Electron paramagnetic
resonance spectroscopy
Determine the reactivity of the dust for biologic implications
5.0
5.0
N/A Static experiment
A few independent measurements
Langmuir probe
Separate regolith particles into >100 nm and <100 nm size
fractions for EPR experiment
Levitated dust
3.0
5.0
Continuous
Particle counter
Levitated dust
7.0
7.5
Continuous
Arm
13.0
43.0
As required for sampling.
Scoop
Deploy inst, conduct experiments, collect samples
Recover surface regolith samples
0.5
0.0
As required for sampling.
Geotechnical Expts
End effector for geotech properties
3.0
0.0
Imaging lidar
Topography of landing region and upper crater interior
13.0
30.0
Scan of crater interior
UV imaging
5.0
5.0
Periodic obs crater interior
Emission spectroscopy
View the interior of the crater with Lyman a illumination
Chemical comp from micrometeorite impact flashes
3.0
7.0
Cont Obs of crater interior
IR Bolometer
Determine surface and near surface temperatures
2.0
5.0
Periodic obs of crater interior
Sample processing for EPRS
As required for EPRS measurment
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Telescope/Reflector
Schafer SLMS (Silicon Lightweight Mirror)
•Communication Hybrid Optical RF Dish
(CHORD)
•40 Cm Dia Primary Mirror, 60 Cm RF
Reflector (12cm Flexible Extensions)
•Weight: .6kg For Substrate + .4kg
Boom + .1kg Horn
•TRL 6 Globalhawk Mirror
Antenna Feed
RF: Prime
Focus
Dichroic reflector
CMOS imager assembly
Optical: Cassegrain Focus
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Cis Lunar Payload
LSAS – Composition of
dust, exosphere, &
surface
ESA, FGM, EFI
Lunar surface potential
UV/Vis sensor –
detect dust
remotely
Dust Analyzer – Q,
v, m of dust grains
Instrument
kg
W
EFI
3.86 0.36
ESA
2.24 1.77
FGM
1.46 0.01
LSAS
3.5
5
IDPU
4.5
7
Imager
2
0.5
Dust analyzer
1
3.5
Reactivity analyzer
2
1
Line scanner
1
0.5
TOTAL 21.56 19.64
DREX – Measures
dust chemical
reactivity
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UNCLASSIFIED//ITAR Restricted
Development
Projects
30 Hz Miniaturized
Polarimeter Minaturized Camera
Component
Dimensions
Weight
VNIR/PI Camera
1.75”x2”x5”
½ lb
Includes lens
Gimbal
5” x 4.4” x 7”
2 lb
Weight with GPS and
electronics
1”x4”x6”
2 lb
Distribution and
conditioning
~5”x5”x12”
5 lb
Power Supply
Communications
Camera and Gimbal
Comments
Dual Transmitters
Onboard Computer
Power Supply
Total (as built)
<15 lbs
Includes structure,
window, etc.
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UNCLASSIFIED//ITAR
Restricted
Micro Lunar Lander Payload Capabilities
•
Notional Capability for 130 kg Lander
•
Payload Mass - 50 Kg max
•
•
dependent on location payload on lander
Payload mass would need to be split between north and
south side of vehicle
•
•
Payload Power
•
•
•
15 Watts continuous, 30 Watts w/50% duty cycle
Short duration peak power < 2 minutes: 50 Watts
Payload Volume
•
•
•
•
Exact split to be dependent on C.G location of each payload
Internally mounted payloads: 7” W x 8”H x 5” D
Externally mounted payloads: 14”W x 10”H x 6” D
Unique payload envelopes such as drills, scoops and robotic
arms would need to be evaluated on a case by case basis
Locations for payload mounting
•
Extension module sidewall panels
•
•
•
Interior and exterior of north facing radiator panel
Interior on south facing solar panel
Upper radiator panel
•
•
Interior as available (shared with avionics)
Exterior (limited by radiator for thermal management)
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Solar Wind Sentinel Instruments
•
Measurement objectives
–
Determination of solar wind composition
•
•
•
ACE instruments
–
–
Principally late-70’s heritage
SIS/SWIMS – solar wind isotope mass spectrometer solar measures high-energy particle flux
•
–
–
Multiple solid-state charged particle detector w/incidence telescope (scanning over sky, apertures ~1
cm2)
SWEPAM – solar wind ions
•
–
–
–
–
Two telescopes followed by stacks of charged particle position-sensitive solid state detectors (aperture
~40 cm2)
EPAM – electron, proton, alpha particles monitor
•
•
Elemental (hydrogen to zinc, Z=1-30), isotopic, and ionic charge state
Energies range from 100 eV to 500 eV
Multiple channels w/collimator, electrostatic analyzer, electron multipliers
MAG – vector magnetometer
ULEIS – ultra-low energy isotope spectrometer
SEPICA – solar energetic particles ionic charge analyzer
CRIS – cosmic-ray isotope spectrometer
State of the art instrument suite would be less than 6 kg / 15 W
–
Based on examples like Swedish Munin spacecraft
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PICO: Primordial Infrared Cosmic Observer
•
•
•
•
•
•
•
Scientific Goal: Detect distant galaxies during the epoch of reionization of the universe at 3.3 and
5 um wavelength. This is near the minimum in the zodiacal background. Goal is to detect objects
to the confusion limit and map a small area if there is remaining mission time to do so. Results will
be significant for understanding initial galaxy formation in the Universe and the nature of first light
objects.
Relation to other Missions: Goal is to go significantly deeper and / or cover greater area than
Spitzer IRAC. 1 yr of PICO should be more sensitive than 1 month of Spitzer. Much more
sensitive than WISE or ASTRO-F since those are survey missions. Might be able to recover some
WISE science if WISE is cancelled. This will be JWST precursor science. Each exposure will have
64x the area of Spitzer IRAC and will have the same size pixels on the sky (~ 1.2”).
Mission Concept:
The mission requires that its instrument be pointed at / near the galactic / ecliptic pole
for about 1 yr duration. The instrument needs to be in a stable thermal environment with few
external heat loads. Geosync may be a possible orbit, a solar drift-away orbit would definitely
work, and it may be possible to site the instrument near 1 of the lunar poles (if the detector can get
cold enough there). If sited on the moon, then the instrument could also function as a site survey
telescope (measure emissivity over time).
The instrument is a very simple 30cm Al telescope with a single off-the-shelf 2k x 2k pixel HAWAII
2RG HgCdTe IR detector array (substrate thinned) with 1 – 5 micron response. 3.3 um (and
possibly 5 um) filters are located just above the detectors. The telescope is passively cooled to
below 70K and the detector is cooled (via a radiator) to below 40K. There is only 1 operating
mode. Communication bandwidth depends on on-board storage and downlink strategy, but is
estimated to be on the order of 1 Mbit / sec .
The spacecraft does need to be 3-axis stabilized if deployed in Geo, solar, or another orbit. RMS
pointing uncertainty needs to be on the order of 1 arcsecond. A lunar lander is required if the
No
Distribution Without Permission
instrument is FOUO
to be sited
onSecondary
the moon.
Space Weather In-situ Hardware
(SWISH) Optimization for the VSE
Mission & Objectives
• NASA needs to place a coherent suite of sensors aboard every
lunar vehicle to measure in-situ and to provide for a standardized
measurement of key parameters of the space radiation environment
spectrum.
• This standardized sensor suite complement will evolve and establish
itself as the "gold standard" by which the same sensors' performance
can be measured repeatedly on every trans-lunar voyage, in lunar
orbit, and eventually on transits to Mars.
• This sensor suite will provide for an instrument validation testbed for
sensors needed by ESMD to support mission objectives such as
astronaut EVA and dosimetry within the manned CEV and lunar
habitat environments.
• Small satellites offer a unique opportunity to mature existing
technologies and evolve new technologies in support of radiation
measurements in space.
Small Satellite TestBed Implementation
• Small sats (100-1000kg) are excellent testbeds since sensors
with their supporting instrumentation can be placed in a variety of
radiation environments (e.g., LEO, highly inclined orbits through
the electron/trapped proton belts, trans-lunar/Martian, lunar orbit,
earth-moon and sun-earth-moon Lagrange points). Example: ST-5
launched March 2006 to inner magnetosphere.
• Small sats allow for several quick iterations to achieve
standardization of a sensor and its supporting architecture.
• Small sats allow for in-situ testing of the sensors in their space
environment for long periods of time (as would be required for
lunar and Martian missions).
Payload Description
• This sensor complement would cover an optimized range of
particle energy, flux, and energy transfer characteristics of
interest to NASA's Vision for Space Exploration.
• It will build upon existing mature radiation sensor instruments
flown aboard work-horse SEC missions such as ACE and SOHO
(e.g., each instrument is relatively low mass (~5-30kg), requires
modest power (few-several 10s of Watts) and telemetry (10s –
1000s bits/s)).
• Lunar Prospector (LP) had three in-situ radiation measurement
instruments smaller in mass, power, and telemetry than the larger
SEC missions.
• The proposed sensor complement can leverage off the recently
launched ST-5 idea of using small satellites with radiation sensor
payload instrumentation.
Cost & Scope
• Development effort is needed to optimize existing high-TRL
sensors suites flown on ACE, SOHO, and LP and validate new
technologies emerging as smaller, less power, and lower
bandwidth radiation sensors are being developed.
• The total LP instrument complement (5 instruments) cost <$3M
(FY94). 3/5 instruments were radiation sensors (e.g., alpha
particle, neutron & X-ray/gamma-ray spectrometers) developed
by LANL.
• The success of ST-5 implies technology exists for reduced
mass & power radiation sensors to be tested, validated and
standardized for future use on missions to the Moon and Mars.
POC: Kimberly Ennico
[email protected]
Tel: 650-604-6067
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Micromagnitude Variability of
Nearby Main Sequence Stars
Mission & Objectives
The ages of the nearby ZAMS stars have not been
determined with precision. Based on the amplitude of their
radial g-mode oscillations in brightness, asteroseismology
offers an interpretive tool for determining the ages of those
stars that are evolving off the main sequence.
The mission is a small telescope in space is able to make
precise observations at the micromagnitude level of
precision, a level not available from ground based
observatories that are limited at the milimagnitude level.
Benefits and Rationale
The theory of stellar evolution predicts the observable
path that will be traced by any given star based on its
initial mass and metallicity. To date, stars at the initial
stages of becoming giants have not been distinguished
from younger ZAMS neighbors. Asteroseismology has
been successful in interpreting millimagnitude amplitude
variability.
An observatory capable of micromagnitude (ppm) stability
and accuracy is not presently available for the brightest
nearby stars. The defunct GP-B fine guidance telescope
has demonstrated the required precision at the 10
micromagnitude level.
Instrument
The telescope is based on the heritage of the flight proven
GP-B fine guidance telescope, thermally stabilized
ultrahigh sensitivity photodetectors, and readout
electronics.
1) 15 cm aperture class telescope having a 2 arcmin field
of view with beam splitters and bandpass filters. 2) Spin
stabilized spacecraft & pointing system with 10 arcsec
pointing capability using microthrusters. 3) The spacecraft
bus will be an available design.
Deliverable & Outcomes
Low cost satellite with spin stabilized pointing system and
a telescope with cryogenic cooler and photometric
detectors for the ultraviolet, visible and infrared.
Determination of the precise ages of stars on the Zero
Age Main Sequence (ZAMS).
Determination of the variability of bright nearby stars
previously not known to be variable at all.
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POC: John Goebel
[email protected]
x 43188
Deuterium Abundance in the Galaxy
Mission & Objectives
Deuterium was formed in the Big Bang, and its abundance
is very sensitive to the conditions at the time it was formed.
Deuterium is easily destroyed in stars, but there are no
known methods for producing it. Thus, its abundance
provides strong constraints on the physical conditions in the
very early universe, and on the subsequent star formation
history of the universe.
Our objective is to measure the deuterium abundance in
PAHs and HDO, two sinks of deuterium, as a function of
star formation activity to determine the destruction rate of
deuterium by stars and the primordial deuterium
abundance.
Benefits and Rationale
Traditional methods using UV lines in absorption to nearby
stars to determine the deuterium abundance show large
variations that can be explained by deuterium depletion
onto dust and molecules. The limited range of the UV
observations cannot address deuterium destruction via
stars. Infrared spectroscopy is well suited for studying the
deuterium abundance in molecules throughout our galaxy
since molecules have their fundamental frequencies in the
infrared, and infrared wavelengths penetrate the dusty
disk of the galaxy.
Instrument
The instrument is a very simple 50cm Al telescope with
a medium spectral resolution (≈1500) echelle
spectrometer using a single off-the-shelf 2k x 2k pixel
HAWAII 2RG HgCdTe IR detector array with 1 – 5 micron
response. The telescope is passively cooled to below 70K
and the detector is cooled (via a radiator) to below 40K.
The instrument needs to be in a stable thermal
environment with few external heat loads; possibly
Geosync, a solar drift-away orbit would definitely work,
and it may be possible to site the instrument near one of
the lunar poles (if the detector can get cold enough there).
Deliverable & Outcomes
Low cost satellite observing system to study the
deuterium abundance as a function of star formation
activity.
Determination of the destruction rate of deuterium.
Determination of the primordial deuterium abundance and
hence the density of baryons in the universe.
POC: Jesse Bregman
[email protected]
x46136
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XNAV Path Forward
• NASA DARPA Partnership
• Shuttle Launch 2010
• ISS Mission Manifested ULF3
• Projects Objectives
•Venture Class Approach
•Navigation, 130 M SEP Anywhere
in Solar System
•X-Ray Astronomy Afforded by
Improved Resolution (3 orders of Mag)
Timing References (6 orders of Mag)
NFOV Sensor &
Electronics
70 FTEs
$8 M
Payload
Support
Processor
Gimbal
Assembly
150 Kg
200W
Atomic
Clock
IMU
GPS
Receiver
GPS Antenna
2PL
2011
PHASE I
Concept Feasibility
Characterize Pulsars
Attitude/position Algorithm
Prototype Detector Design
Prototype Sensor Design
CONOPS Development
PHASE II
GSE Development
BAA
PAD
Signed
CoDR
CDR
PDR
Launch
ULF3
PDR
CoDR
P-II
Go/No-Go
(Re-compete)CDR
Competition / Source Selection
Design Development
Fabrication / Assembly
Space Qualification
GSE Hardware Development
PHASE III
P-III
Go/No-Go
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Phase I
PhasePermission
II
Data
Collection
& Analysis
Phase III
XNAV Payload
Functional Architecture
GSFC
ARC
ARC
GSFC
ARC
GSFC
ARC
ARC
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On-Orbit Anomalies - 2003
*Extracted from Orbital Anomalies in Goddard Spacecraft for Fiscal Year 2003
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NanoSat for Solar Wind Monitoring
•
ACE background
– ACE (Advanced Composition Explorer, launch in 1997) proved to be valuable
asset for near-real-time monitoring of solar wind
– Developed unintended addition to its basic research role by providing significant
operational value of ~one hour advanced warning of geomagnetic storms
– Large spacecraft (~785 kg at Delta-2 launch, early PI-led mission)
– Desire for long-term replacement solution
• ACE exceeding significantly beyond its design lifetime
– Recurring launches with possible redundant system
– Many studies and proposals over past ten years
• Either too expensive or not from credible players
•
Solar Wind Sentinel mission
– Earth-Sun L-1 libration point (unstable)
• ~1.5 million km from Earth, approximately 200,000x50,000 halo orbit
– Propulsion requirements
• LEO injection (requires solid kick stage for Falcon-1 launch, slightly more dv than lunar
mission, +35 m/sec)
• L-1 halo orbit capture (<50 m/sec)
• Moderate halo orbit maintenance (~10 m/sec/year)
• Reaction control (minimal if solar radiation pressure can be managed)
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Small Sat Investments $M
DARPA+NRO
Existing
In Orbit
Planned
Total
1000
450
750
2200
135
427
562
1917
1917
150
321
140
140
60
280
340
230
912
1192
250
5250
4826
11922
ORS
AF AIRSS
SSTL
171
RAPIDEYE
ESA
CNES
50
MDA
5000
TOTAL
6050
1046
30+smallsats
DARPA/NRO: F6, M idstep, Spawn, Roast, M itex, Streak, Isat, Dsx, Glomr, Tercel/Secs, M acsat1,2, M icrosat1-7, Darpasat, other
ORS: XSS-11, Tacsat 1,2,3, ORS PM D
SSTL: Uosat
1-5,S80/T,Kitsat,Posat,Healthsat,Fasat
Small
Sat
Investments in AB,Clementine,Ceris,Thai-Paht,Tiunsat,Snap,Tsinghua,Picosat,Alsat,Bilsat,Nigeria,UK-dmc,Topsat,Beijing
Billions: Yesterday, Today, Tomorrow
ESA: PROBA1,2,3,4
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Permission
CNES: Demeter,
Parasol, Essaim(4),
Spirale(2), M icroscope, Picard,
HRG (4), HRG+, GMWithout
ES, Pegase, Taranis,
SM ES, Altika, ALsat2
Small Sat Cost, Weight, Performance
Bus $M
14.7
12.8
17.8
30.0
20.0
24.0
6.0
12.0
25.0
<20
<20
<20
lo cost
lo cost
22.0
20.4
20.0
15.7
15.7
16.0
SC Kg Bus Kg PL Kg
149.0
99
50
360.0
200
160
450.0
400
50
142.0
117
25
142.0
117
25
240.0
200
40
67.0
53
14
157.0
99
58
170.0
85
85
110.0
66
44
166.0
115
41
150.0
115
35
94.0
39
55
92.0
62
30
135.0
85
50
135.0
85
50
120.0
75
45
110.0
60
50
150.0
80
70
120.0
80
40
Swales
Swales
Swales
LM
LM
BAC
Dynacon
Aero Astro
Ball
BNSC
SSTL
MDA
ESA
DLR
CNES
CNES
CNES
CNES
CNES
CNES
Themis
ORS
ORS+
XSS11
XSS11b
ORS
MOST
STPsat
STP/SIV
Topsat
Beijing-1
Rapid Eye
Proba
BIRD
Demeter
Parasol
Essaim
Myriade1
Myriade2
Spirale
NASA
NASA
NASA
LLO
LO
LL
references
1. Demeter is up and running, Space New 24 August 2004
27.0
25.0
25.0
158.0
200.0
130.0
134
122
81
24
78
49
PL % DV m/s
34
950
44
0
11
950
18
650
18
650
17
900
21
0
37
0
50
0
40
0
26
0
23
0
59
0
33
0
37
80
37
80
38
90
45
100
47
75
33
90
15
39
38
491
184
609
Stabil. Power W
Spin
40
3 Axis
600
3 Axis
600
3 Axis
380
3 Axis
380
3 Axis
650
3 Axis
35
3 Axis
160
3 Axis
200
3 Axis
40
3 Axis
40
3 Axis
40
3 Axis
40
3 Axis
60
3 Axis
200
3 Axis
200
3 Axis
200
3 Axis
200
3 Axis
200
3 Axis
200
3 Axis
3 Axis
3 Axis
FFP quote
Post CDR
eng est.
Cost from AFRL CDRLs
quote
Proposal
See Ref 6
See Ref 7
Post CDR
$38.9 Mission inc PL, ground, launch, refs 8,9
$29.2 Mission inc PL, ground, launch, ref 8,
$140 FFP: 5 spacecraft inc PL, ground systems, refs 8, 10,
$52M mission inc PL, rideshare, operations, devt
15.6M euro includes payload ref 4
$123M mission cost, 3 years of operatiosn, ref 2
ref 1 for cost, ref 3 for weights
ref 1 for cost, ref 3 for weights
$162 mission, inc launch ops payloads, 2 microsats, ref 5
231
231
153
2. CNES readies Essaim Satellites for December Launch, Space News 11 October 2004
3. CNES micro sat program, 20th annual conference on small satellites, 14 Aug 2006, USU
4. Parasol a microsat in the A-train for earth atmospheric observations IAA Conf 2005
5. Press Release French Embassy Washington DC.28 Jan 2004
6. The Most microsat mission: 1 year in orbit, 18th annual conference on small satellites, 14 Aug 2006, USU, SSC 04 IX-1
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7. STPSat-1 Capabilities and Overview, Aero Astro
NASA SMEX Heritage
FAST – 8/96
SAMPEX – 7/92
Study solar, anomalous, galactic, and
magnetospheric energetic particles
36 months, $53M development
· S/C 258 lbs, 60 watts
· P/L 88 lbs, 22 watts
· Zenith oriented sun pointer
SWAS – 12/98
Plasma physics investigation of high
altitude aurora
42 months, $45M development
· S/C 284 lbs, 33 watts
· P/L 112 lbs, 15 watts
· Spin stabilized, magnetically processed
Investigation into the composition
of dense interstellar clouds
60 months, $64M development
· S/C 410 lbs, 133 watts
· P/L 225 lbs, 59 watts
· 3-axis stabilized, fine stellar pointer
TRACE – 4/98
WIRE – 3/99
Explore and define the dynamics and
structure of the solar heliosphere
Survey starburst galaxies in the far-infrared
to determine their evolutionary rates
36 months, $40M development
· S/C 348 lbs, 114 watts
· P/L 97 lbs, 30 watts
· 3 –axis stabilized, fine sun pointer
46 months, $46M development
· S/C 403 lbs, 125 watts
· P/L 154 lbs, 34 watts
· 3 –axis stabilized, fine sun pointer
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Lunar Express Orbiter
Patch Antennas
Star Tracker
Lasercom
Battery
Transmitter
Reaction Wheel
Receiver
Amplifier
IPP Router
& Local RF Comm
Leverage Flight Heritage
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31
Common Bus Block Diagram
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