Pete Klupar Peter.D.Klupar@nasa

1. Advantages of Very Small Spacecraft 15 May, 2007 Pete Klupar Peter.D.Klupar@nasa.gov

2. Definitions Development Mass Cost Time Large 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 First Proposed By Surrey Satellite Technology Limited

3. 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

4. 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)

5. Background – International Activities International Efforts Include >1000 Small Satellites

6. Emerging Small ELVs Offer Cost Effective Performance Unique opportunity for increased mass at substantially lower cost

7. 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 • Stage 1 -SR118 • Performance: • 496 Kg to TLI • Total Launch Cost (ROM): • ~$36M (First Mission) • Includes S-37GV Qual • ~$26M (Recurring)

9. 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

10. Notional Costs and Schedule $10M Budgetary Cost ($) $5M 6 12 18 Payload Delivery or Availability Schedule (months) NFIRE Optimized Design Based on existing Instrument MITEC SPARE with New 7.5cm Telescope MSO SPARE with New 30cm Telescope MSTI-3 SPARE MISTEC with 16cm Telescope & New FPA SIS/SWIMS SPARE EPAM(as is) SWEPAMwith upgraded sensor

11. 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

12. Small Concept Star Tracker Diplexer Transmitter Patch Antennas Receiver Amplifier Avionics Additional payload space as available Battery North side panel for externally mounted payloads DSMAC Radar Altimeter Payload(s) located internally

13. Small Lander Payloads

14. Telescope/Reflector RF: Prime Focus Antenna Feed Dichroic reflector CMOS imager assembly Optical: Cassegrain Focus 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

15. Cis Lunar Payload LSAS – Composition of dust, exosphere, & surface ESA, FGM, EFILunar surface potential DREX – Measures dust chemical reactivity UV/Vis sensor – detect dust remotely Dust Analyzer – Q, v, m of dust grains

16. Development Projects UNCLASSIFIED//ITAR Restricted Camera and Gimbal 30 Hz Miniaturized Polarimeter Minaturized Camera Dual Transmitters Onboard Computer Power Supply UNCLASSIFIED//ITAR Restricted

17. 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)

18. 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 cm2) • EPAM – electron, proton, alpha particles monitor • Multiple solid-state charged particle detector w/incidence telescope (scanning over sky, apertures ~1 cm2) • 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

19. 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 instrument is to be sited on the moon.

20. Space Weather In-situ Hardware(SWISH) Optimization for the VSE Mission & Objectives Payload Description • 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. • 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 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. • 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). POC: Kimberly EnnicoKimberly.A.Ennico@nasa.govTel: 650-604-6067

21. Micromagnitude Variability of Nearby Main Sequence Stars Mission & Objectives 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. 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 Deliverable & Outcomes 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. 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 jjgoebel@arc.nasa.gov x 43188

22. Deuterium Abundance in the Galaxy Mission & Objectives Instrument 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. 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). Benefits and Rationale 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. 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. POC: Jesse Bregman Jesse.Bregman@nasa.gov x46136

24. XNAV Path Forward 2011 • 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 Launch ULF3 BAA CDR CoDR PDR PAD Signed PDR CoDR P-II Go/No-Go (Re-compete) CDR P-III Go/No-Go Data Collection & Analysis Phase I Phase III Phase II • 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

25. XNAV PayloadFunctional Architecture GSFC ARC ARC GSFC ARC GSFC ARC ARC

26. On-Orbit Anomalies - 2003 *Extracted from Orbital Anomalies in Goddard Spacecraft for Fiscal Year 2003

27. 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)

28. Small Sat Investments $M Small Sat Investments in Billions: Yesterday, Today, Tomorrow

29. Small Sat Cost, Weight, Performance

30. NASA SMEX Heritage FAST – 8/96 Plasma physics investigation of high altitude aurora SWAS – 12/98 Investigation into the composition of dense interstellar clouds 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 42 months, $45M development · S/C 284 lbs, 33 watts · P/L 112 lbs, 15 watts · Spin stabilized, magnetically processed 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 Explore and define the dynamics and structure of the solar heliosphere WIRE – 3/99 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

31. Lunar Express Orbiter Patch Antennas Star Tracker Lasercom Battery Transmitter Reaction Wheel Receiver Amplifier IPP Router & Local RF Comm Leverage Flight Heritage

32. Common Bus Block Diagram