pete klupar peter.d.klupar@nasa

Pete Klupar

[email protected]

Advantages of Very Small Spacecraft

15 May, 2007

FOUO No Secondary Distribution Without Permission

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


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


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)


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


Emerging Small ELVs Offer Cost Effective Performance

Unique opportunity

for increased mass at

substantially lower cost

Launch Vehicle

200 km, 38 °

LEO Mass (kg)

Estimated

GTO Mass (kg)

EstimatedTLI Mass (kg)

Estimated

NEO Mass (kg)

EstimatedFairing Diameter

(m)Price ROM ($M)

Pegasus 425 N/A N/A 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 N/A N/A 1.3 $25

Minotaur 4/5 1700 692 464 425 2.3 $25 - $31

Falcon 1 570 107 82 N/A 1.5 $10 - $12


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)


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


Notional Costs and Schedule

Payload Delivery or Availability Schedule (months)

126 18

$10M

$5M

Budgetary Cost ($)

EPAM(as is)

SWEPAM with upgraded sensor

Optimized Design Based on existing Instrument

MSTI-3 SPARE

MISTEC with 16cm Telescope & New FPA

MSO SPARE with New 30cm Telescope

NFIRE

MITEC

SPARE with New 7.5cm Telescope

SIS/SWIMS SPARE


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


Small Concept

PatchAntennas

Star Tracker Diplexer

RadarAltimeter

Avionics

Transmitter

AmplifierReceiver

Battery

DSMACPayload(s) located internally

Additional payload spaceas available

North side panel for externally mounted payloads


Lander Payload Element ObjectiveMass (kg)

Power (W)

Duty Cycle

Stereo imaging system Surface images for analysis 0.8 6.0 360° images

Mast for stereo imager Provide elevation for imaging 3.5 9.5 1 deployment

Drill, deployment mech samples from depths of 2 m 20.0 30.0 2-4 hrs at station

Gas Chromatograph Mass Spectrometer

Determine volatile compounds and isotopic composition 19.0 75.02-hour analysis measurement

Sample processing system for GCMS

Process core or scoop material for analysis.For each GCMS sample

Beacon Navigation reference 1.0 5.0

Magnets Magnetic susceptibility of regolith particles 0.5 0.0 N/A Static experiment

Electron paramagnetic resonance spectroscopy

Determine the reactivity of the dust for biologic implications5.0 5.0

A few independent measurements

Sample processing for EPRSSeparate regolith particles into >100 nm and <100 nm size fractions for EPR experiment

As required for EPRS measurment

Langmuir probe Levitated dust 3.0 5.0 Continuous

Particle counter Levitated dust 7.0 7.5 Continuous

Arm Deploy inst, conduct experiments, collect samples 13.0 43.0 As required for sampling.

Scoop 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 View the interior of the crater with Lyman a illumination 5.0 5.0 Periodic obs crater interior

Emission spectroscopy 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

Small Lander Payloads


Telescope/Reflector

•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

Schafer SLMS (Silicon Lightweight Mirror)

Dichroic reflector

Antenna Feed

CMOS imager assembly

Optical: Cassegrain Focus

RF: Prime Focus


Cis Lunar Payload

Dust Analyzer – Q, v, m of dust grains

ESA, FGM, EFILunar surface potential

LSAS – Composition of dust, exosphere, &

surface

UV/Vis sensor – detect dust remotely

DREX – Measures dust chemical

reactivity

Instrument kg WEFI 3.86 0.36ESA 2.24 1.77FGM 1.46 0.01LSAS 3.5 5IDPU 4.5 7Imager 2 0.5Dust analyzer 1 3.5Reactivity analyzer 2 1Line scanner 1 0.5

TOTAL 21.56 19.64


Development Projects

30 Hz Miniaturized Polarimeter Minaturized Camera

UNCLASSIFIED//ITAR Restricted

UNCLASSIFIED//ITAR Restricted

Dual Transmitters Onboard Computer

Power Supply

Camera and Gimbal

Includes structure, window, etc.

<15 lbsTotal (as built)

5 lb~5”x5”x12”Communications

Distribution and conditioning

2 lb1”x4”x6”Power Supply

Weight with GPS and electronics

2 lb5” x 4.4” x 7”Gimbal

Includes lens½ lb1.75”x2”x5”VNIR/PI Camera

CommentsWeightDimensionsComponent

Includes structure, window, etc.

<15 lbsTotal (as built)

5 lb~5”x5”x12”Communications

Distribution and conditioning

2 lb1”x4”x6”Power Supply

Weight with GPS and electronics

2 lb5” x 4.4” x 7”Gimbal

Includes lens½ lb1.75”x2”x5”VNIR/PI Camera

CommentsWeightDimensionsComponent


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)


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


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.


Space Weather In-situ Hardware(SWISH) Optimization for the VSE

Mission & Objectives

Small Satellite TestBed Implementation

Payload Description

Cost & Scope

• 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 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).

• 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 [email protected]: 650-604-6067


Micromagnitude Variability of Nearby Main Sequence Stars


Benefits and Rationale

Instrument

Deliverable & OutcomesThe 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.

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.

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.

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 [email protected] x 43188


Deuterium Abundance in the Galaxy


Benefits and Rationale

Instrument

Deliverable & OutcomesTraditional 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.

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).

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 [email protected] x46136


XNAV Path ForwardXNAV Path Forward

GPS Antenna2PL

Atomic Clock

IMU

NFOV Sensor & Electronics

Gimbal Assembly

PayloadSupport

Processor

GPSReceiver

150 Kg200W

• NASA DARPA PartnershipNASA DARPA Partnership• Shuttle Launch 2010Shuttle Launch 2010• ISS Mission Manifested ULF3ISS Mission Manifested ULF3• Projects ObjectivesProjects Objectives

•Venture Class ApproachVenture Class Approach•Navigation, 130 M SEP Anywhere Navigation, 130 M SEP Anywhere in Solar Systemin Solar System•X-Ray Astronomy Afforded by X-Ray Astronomy Afforded by Improved Resolution (3 orders of Mag) Improved Resolution (3 orders of Mag)

Timing References (6 orders of Mag) Timing References (6 orders of Mag)

PHASE I

Concept FeasibilityCharacterize Pulsars

Attitude/position Algorithm

Prototype Detector Design

Prototype Sensor Design

CONOPS Development

PHASE II

GSE DevelopmentCompetition / Source Selection

Design Development

Fabrication / Assembly

Space Qualification

GSE Hardware Development

PHASE III

BAA PADSigned

CoDR PDR CDR

P-IIGo/No-Go

(Re-compete)CDR

Data Collection& Analysis

P-IIIGo/No-Go

CoDR

PDR

Phase I Phase II Phase III

LaunchULF3

2011

70 FTEs$8 M


XNAV PayloadFunctional Architecture

ARC

GSFC

ARC

ARC

ARC

GSFC

GSFC

ARC


On-Orbit Anomalies - 2003

*Extracted from Orbital Anomalies in Goddard Spacecraft for Fiscal Year 2003


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)


Small Sat Investments $M

Existing In Orbit Planned Total

DARPA+NRO 1000 450 750 2200

ORS 135 427 562

AF AIRSS 1917 1917

SSTL 171 150 321

RAPIDEYE 140 140

ESA 60 280 340

CNES 50 230 912 1192

MDA 5000 250 5250

TOTAL 6050 1046 4826 11922

30+smallsats

DARPA/NRO: F6, M idstep, Spawn, Roast, M itex, Streak, Isat, Dsx, Glomr, Tercel/Secs, Macsat1,2, M icrosat1-7, Darpasat, other

ORS: XSS-11, Tacsat 1,2,3, ORS PMD

SSTL: Uosat 1-5,S80/T,Kitsat,Posat,Healthsat,Fasat AB,Clementine,Ceris,Thai-Paht,Tiunsat,Snap,Tsinghua,Picosat,Alsat,Bilsat,Nigeria,UK-dmc,Topsat,Beijing

ESA: PROBA1,2,3,4

CNES: Demeter, Parasol, Essaim(4), Spirale(2), M icroscope, Picard, HRG (4), HRG+, GMES, Pegase, Taranis, SMES, Altika, ALsat2

MDA: LOSAT, STRV, MSTI, Clementine, LEAP, EKV, THAAD, ASAT, SBI, BP, Intellect, KKVWS, Have Sting, Gremlin, NFIRE

Small Sat Investments in Billions: Yesterday, Today, Tomorrow


Small Sat Cost, Weight, Performance


NASA SMEX Heritage

36 months, $53M development· S/C 258 lbs, 60 watts· P/L 88 lbs, 22 watts

· Zenith oriented sun pointer

SAMPEX – 7/92Study solar, anomalous, galactic, and

magnetospheric energetic particles


· Spin stabilized, magnetically processed

FAST – 8/96

Plasma physics investigation of high altitude aurora


· 3-axis stabilized, fine stellar pointer

SWAS – 12/98

Investigation into the composition of dense interstellar clouds

36 months, $40M development· S/C 348 lbs, 114 watts

· P/L 97 lbs, 30 watts· 3 –axis stabilized, fine sun pointer

TRACE – 4/98

Explore and define the dynamics and structure of the solar heliosphere


· 3 –axis stabilized, fine sun pointer

WIRE – 3/99

Survey starburst galaxies in the far-infrared to determine their evolutionary rates

31


Lunar Express Orbiter

BatteryTransmitter

Receiver

Amplifier

Reaction Wheel

Patch Antennas Star Tracker

Lasercom

IPP Router

& Local RF Comm

Leverage Flight Heritage


Common Bus Block Diagram

pete klupar peter.d.klupar@nasa

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