Pete Klupar [email protected]

Advantages of Very Small Spacecraft 15 May, 2007

Definitions

Large Development Mass 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

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

United Kingdom ESA France Japan Sweden Germany Denmark Israel Canada India Others

Small Satellite Programs

SSTL ~ 40 missions <$100M, 5-500Kg; DERA/QINETIQ (STRV) Smart-1, PROBA-1, PROBA 2……PROBA-N CNES - Myriade <150kg S/C, <70kg P/L, 6 launched since 2004, 10 in development JAXSA – Index (72 Kg, 2005 launch <$10M) Swedish Space Corp – 6 Small/Microsats in orbit, 3+ in development (Viking, Freja, Astrid 1,2 Odin, Prisma, Svea etc) DLR, TuB (TUBSAT-A, -B, -N/N1,-DLR, -MAROC,- LAPAN) DTU, Terma – Oerstad, Romer Rafael, IAI – EROS-A, EROS-B (Imaging Microsatellites) Dynacon/UTIAS – MOST, NESS, Brite, MDA – Rapid Eye ISRO – HAMSat (45 kg microsatellite) 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 Taurus 3110/3113 Taurus 3210

200 km, 38 ° LEO Mass (kg)

Estimated

425 1530 1291

GTO Mass (kg)

N/A

Estimated

627 107

TLI Mass (kg)

Estimated

N/A 427 367

NEO Mass (kg)

Estimated

N/A ?

N/A

Fairing Diameter (m) Price ROM ($M)

1.3

1.6

2.3

$35 $50 $50 Minotaur 1 Minotaur 4/5 Falcon 1 Unique opportunity for increased mass at substantially lower cost 565 1700 570 N/A 692 107 N/A 464 82 N/A 425 N/A 1.3

2.3

1.5

$25 $25 - $31 $10 - $12 FOUO No Secondary Distribution Without Permission

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 – Only Subset Required to Fly Stage 5 Cold Gas Attitude Control System (ACS) Composite Structure • • • • • • •

Guidance Control Assembly (GCA)/Stage 4

GCA Design Shared with Minotaur III & IV OSP-Standard Flight Proven Avionics – Split Between S4 and S5 • Cold Gas ACS Stage 4 Star 48V SRM (New for M-V) – Thrust Vector Control – Qualified via Static Fire •

GFE Peacekeeper Stages

Stage 3 - SR120 Stage 2 - SR119

Performance:

– 496 Kg to TLI

Total Launch Cost (ROM):

– ~$36M (First Mission) • Includes S-37GV Qual – ~$26M (Recurring)

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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

Optimized Design Based on existing Instrument

$10M

MITEC SPARE with New 7.5cm Telescope

Budgetary Cost ($) $5M

EPAM(as is) SIS/SWIMS SPARE MSTI-3 SPARE MSO SPARE with New 30cm Telescope MISTEC with 16cm Telescope & New FPA SWEPAM with upgraded sensor NFIRE

6 12 Payload Delivery or Availability Schedule 18

Recurring Cost ROM

• Overall Recurring Goal For 5

th

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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Patch Antennas Avionics Additional payload space as available

Small Concept

Star Tracker Diplexer Transmitter Receiver Amplifier Battery

North side panel for externally mounted payloads

Radar Altimeter Payload(s) located internally FOUO No Secondary Distribution Without Permission DSMAC

Small Lander Payloads

Lander Payload Element Objective

Stereo imaging system Mast for stereo imager Drill, deployment mech Gas Chromatograph Mass Spectrometer Sample processing system for GCMS Beacon Magnets Electron paramagnetic resonance spectroscopy Surface images for analysis Provide elevation for imaging samples from depths of 2 m Determine volatile compounds and isotopic composition Process core or scoop material for analysis.

Navigation reference Magnetic susceptibility of regolith particles Determine the reactivity of the dust for biologic implications Sample processing for EPRS Langmuir probe Particle counter Arm Scoop Geotechnical Expts Imaging lidar UV imaging Emission spectroscopy IR Bolometer Separate regolith particles into >100 nm and <100 nm size fractions for EPR experiment Levitated dust Levitated dust Deploy inst, conduct experiments, collect samples Recover surface regolith samples End effector for geotech properties Topography of landing region and upper crater interior View the interior of the crater with Lyman a illumination Chemical comp from micrometeorite impact flashes Determine surface and near surface temperatures

(kg)

0.8

3.5

20.0

19.0

(W)

6.0

9.5

30.0

Duty Cycle

360° images 1 deployment 2-4 hrs at station 2-hour analysis measurement 75.0

For each GCMS sample 1.0

0.5

5.0

5.0

0.0

5.0

N/A Static experiment A few independent measurements As required for EPRS measurment 3.0

7.0

13.0

0.5

3.0

13.0

5.0

3.0

2.0

5.0

7.5

43.0

0.0

0.0

30.0

5.0

7.0

5.0

Continuous Continuous As required for sampling.

As required for sampling.

Scan of crater interior Periodic obs crater interior Cont Obs of crater interior Periodic obs of crater interior FOUO No Secondary Distribution Without Permission

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

ESA, FGM, EFI Lunar surface potential UV/Vis sensor detect dust remotely – LSAS – Composition of dust, exosphere, & surface DREX – Measures dust chemical reactivity Instrument

EFI ESA FGM LSAS IDPU Imager Dust analyzer

kg

3.86

2.24

1.46

3.5

4.5

2 1

W

0.36

1.77

0.01

5 7 0.5

3.5

Reactivity analyzer Line scanner 2 1 1 0.5

TOTAL 21.56 19.64

Dust Analyzer – Q, v, m of dust grains

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UNCLASSIFIED//ITAR Restricted

Development Projects

30 Hz Miniaturized Polarimeter Minaturized Camera

Camera and Gimbal Dual Transmitters 1.75

”x2”x5” 5 ” x 4.4” x 7” 1 ”x4”x6” ~5 ”x5”x12” Distribution and conditioning Power Supply Onboard Computer

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 Exact split to be dependent on C.G location of each payload Payload Power • • 15 Watts continuous, 30 Watts w/50% duty cycle Short duration peak power < 2 minutes: 50 Watts • • • Payload Volume 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) FOUO No Secondary Distribution Without Permission

Solar Wind Sentinel

Instruments

• • • Measurement objectives – Determination of solar wind composition • Elemental (hydrogen to zinc, Z=1-30), isotopic, and ionic charge state • Energies range from 100 eV to 500 eV ACE instruments – Principally late-70’s heritage – – – –

SIS/SWIMS

– solar wind isotope mass spectrometer solar measures high-energy particle flux • Two telescopes followed by stacks of charged particle position-sensitive solid state detectors (aperture ~40 cm 2 )

EPAM

• – electron, proton, alpha particles monitor Multiple solid-state charged particle detector w/incidence telescope (scanning over sky, apertures ~1 cm 2 )

SWEPAM

– solar wind ions • 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 FOUO No Secondary Distribution Without Permission

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 FOUO No Secondary Distribution Without Permission

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.

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.

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).

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]

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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. POC: John Goebel [email protected]

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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

a medium spectral resolution (≈1500) echelle with 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]

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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)

Payload Support Processor Atomic Clock IMU GPS Antenna 2PL PHASE I Concept Feasibility Characterize Pulsars Attitude/position Algorithm Prototype Detector Design Prototype Sensor Design CONOPS Development PHASE II GSE Development Competition / Source Selection Design Development Fabrication / Assembly Space Qualification GSE Hardware Development PHASE III BAA PAD Signed CoDR CoDR PDR PDR CDR P-II Go/No-Go (Re-compete)CDR Launch ULF3 NFOV Sensor & Electronics

70 FTEs $8 M

Gimbal Assembly

150 Kg 200W

GPS Receiver 2011 P-III Go/No-Go Data Collection & Analysis

Phase III

ARC

XNAV Payload

Functional Architecture

GSFC GSFC ARC GSFC ARC ARC ARC

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On-Orbit Anomalies - 2003

*Extracted from Orbital Anomalies in Goddard Spacecraft for Fiscal Year 2003 FOUO No Secondary Distribution Without Permission

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) FOUO No Secondary Distribution Without Permission

Small Sat Investments $M

DARPA+NRO ORS AF AIRSS SSTL RAPIDEYE ESA CNES MDA

TOTAL

Existing 1000 50 5000

6050

In Orbit 450 135 171 60 230 Planned 750 427 1917 150 140 280 912 250

4826 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 ESA: PROBA1,2,3,4 FOUO No Secondary Distribution Without Permission CNES: Demeter, Parasol, Essaim(4), Spirale(2), M icroscope, Picard, HRG (4), HRG+, GM ES, Pegase, Taranis, SM ES, Altika, ALsat2 M DA: LOSAT, STRV, M STI, Clementine, LEAP, EKV, THAAD, ASAT, SBI, BP, Intellect, KKVWS, Have Sting, Gremlin, NFIRE Total 2200 562 1917 321 140 340 1192 5250

11922

Small Sat Cost, Weight, Performance

Swales Swales Swales LM LM BAC Dynacon Aero Astro Ball BNSC SSTL MDA ESA DLR CNES CNES CNES CNES CNES CNES NASA NASA NASA Themis ORS ORS+ XSS11 XSS11b ORS MOST STPsat STP/SIV Topsat Beijing-1 Rapid Eye Proba BIRD Demeter Parasol Essaim Myriade1 Myriade2 Spirale 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

LLO LO LL 27.0

25.0

25.0

SC Kg 149.0

360.0

450.0

142.0

142.0

240.0

67.0

157.0

170.0

110.0

166.0

150.0

94.0

92.0

135.0

135.0

120.0

110.0

150.0

120.0

158.0

200.0

130.0

Bus Kg PL Kg 99 200 400 50 160 50 117 117 200 53 99 25 25 40 14 58 85 66 115 115 39 62 85 85 75 60 80 80 85 44 41 35 55 30 50 50 45 50 70 40 134 122 81 24 78 49 50 40 26 23 59 33 37 37 38 45 47 33 PL % DV m/s 34 44 11 950 0 950 18 18 17 21 37 650 650 900 0 0 0 0 0 0 0 0 80 80 90 100 75 90 Stabil. Power W Spin 3 Axis 3 Axis 40 600 600 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 380 380 650 35 160 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 3 Axis 200 40 40 40 40 60 200 200 200 200 200 200 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, 11 $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 15 39 38 491 184 609 3 Axis 3 Axis 3 Axis 231 231 153 references 1. Demeter is up and running, Space New 24 August 2004 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 FOUO No Secondary Distribution Without Permission 8. Microsats - helping to improve security from space, Sweeting, Collective security in space, 15 May 2006 9. Topsat - High resolution imaging from a small satellite, Brooks, Qinetiq, Small Sat Conference USU, SSC01-I-2 10. Rapid Eye an earth observation smallsat constellation, MDA, SSTL, Small Sat Conf, USU, Aug 2003 11. Changing the value proposition of operational space missions, Responsive Space 2005

SAMPEX – 7/92

Study solar, anomalous, galactic, and magnetospheric energetic particles

NASA SMEX Heritage

FAST – 8/96

Plasma physics investigation of high altitude aurora

SWAS – 12/98

Investigation into the composition of dense interstellar clouds 36 months, $53M development · S/C 258 lbs, 60 watts · P/L 88 lbs, 22 watts · Zenith oriented sun pointer

TRACE – 4/98

42 months, $45M development · S/C 284 lbs, 33 watts · P/L 112 lbs, 15 watts · Spin stabilized, magnetically processed

WIRE – 3/99

60 months, $64M development · S/C 410 lbs, 133 watts · P/L 225 lbs, 59 watts · 3-axis stabilized, fine stellar pointer 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 46 months, $46M development · S/C 403 lbs, 125 watts · P/L 154 lbs, 34 watts FOUO No Secondary Distribution Without Permission

Patch Antennas

Lunar Express Orbiter

Star Tracker Lasercom Battery Transmitter Receiver Amplifier IPP Router & Local RF Comm

Leverage Flight Heritage

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Reaction Wheel

31

Common Bus Block Diagram

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