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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 FOUO No Secondary DistributionFirst 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 FOUO No Secondary Distribution Without Permission 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) FOUO No Secondary Distribution Without Permission 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 <$10M) Sweden Swedish Space Corp – 6 Small/Microsats in orbit, 3+ in development (Viking, Freja, Astrid 1,2 Odin, Prisma, Svea etc) Germany DLR, TuB (TUBSAT-A, -B, -N/N1,-DLR, -MAROC,- LAPAN) Denmark DTU, Terma – Oerstad, Romer Israel Rafael, IAI – EROS-A, EROS-B (Imaging Microsatellites) Canada Dynacon/UTIAS – MOST, NESS, Brite, MDA – Rapid Eye India ISRO – HAMSat (45 kg microsatellite) Others China, South Africa, Turkey, Chile, Nigeria, Korea, Taiwan, Australia, Eqypt, Indonesia, Russia, Malaysia, Belgium International Efforts Include >1000 Small Satellites FOUO No Secondary Distribution Without Permission Emerging Small ELVs Offer Cost Effective Performance Launch Vehicle Pegasus 200 km, 38 ° LEO Mass (kg) Estimated 425 GTO Mass (kg) Estimated N/A TLI Mass (kg) Estimated N/A NEO Mass (kg) Estimated Fairing Diameter (m) Price ROM ($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 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 • • 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 FOUO No Secondary Distribution Without Permission FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Permission 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) FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Permission 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 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 FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Permission 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. FOUO No Secondary Distribution Without Permission 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) FOUO No Secondary Distribution Without Permission 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 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 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 FOUO No Secondary Distribution Without Permission 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. FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Permission FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Phase I PhasePermission II Data Collection & Analysis Phase III XNAV Payload Functional Architecture GSFC ARC ARC GSFC ARC GSFC ARC ARC FOUO No Secondary Distribution Without Permission 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 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 FOUO No Secondary Distribution 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 FOUO No Secondary Distribution Without Permission 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 FOUO No Secondary Distribution Without Permission Lunar Express Orbiter Patch Antennas Star Tracker Lasercom Battery Transmitter Reaction Wheel Receiver Amplifier IPP Router & Local RF Comm Leverage Flight Heritage FOUO No Secondary Distribution Without Permission 31 Common Bus Block Diagram FOUO No Secondary Distribution Without Permission