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IAF SYMPOSIUM, REDONDO BEACH, CA, APRIL 19-21, 1999

University Nanosatellite ProgramMr. Maurice Martin, AFRL, martinr@plk.af.mil, (505) 853-4118, Dr. Howard Schlossberg, AFOSR,

Mr. Joe Mitola, DARPA, Dr. Dave Weidow, GSFC, Mr. Andrew Peffer, AFRL, Dr. Richard Blomquist, CMU,Dr. Mark Campbell, UW, Dr. Christopher Hall, VT, Elaine Hansen, UCol, Dr. Stephen Horan, NMSU, Prof Chris Kitts,

Santa Clara Univ, Dr. Frank Redd, USU, Dr. Helen Reed, ASU, Dr. Harlan Spence, BU, Prof Bob Twiggs, Stanford

Abstract

DoD, NASA, and industry are jointly sponsoring thedevelopment and launch of 10 university nanosatellites(~10kg) to demonstrate miniature bus technologies,formation flying, and distributed satellite capabilities.The satellites are planned to launch late 2001. Thispaper provides detailed information on the technologiesand science objectives of each of the 10 satellites.

Program Overview

The Air Force Office of Scientific Research (AFOSR)and the Defense Advanced Research Projects Agency(DARPA) are jointly funding 10 universities with grantsof $50k/year over two years to design and assemble 10nanosatellites (~10kg). The universities will conductcreative low-cost space experiments to explore themilitary usefulness of nanosatellites in such areas asformation flying, enhanced communications,miniaturized sensors, attitude control, and maneuvering.

The Air Force Research Laboratory (AFRL) isdeveloping a deployment structure, securing a launch,and providing such advanced microsatellite hardware ashigh efficiency solar cells and micropropulsion units.NASA Goddard has also teamed with the universities toprovide approximately $1.2M funding to demonstratesuch formation flying technologies as advanced crosslinkcommunication and navigation hardware and flightcontrol algorithms. Numerous industry partners are alsosupporting the universities with hardware and design andtesting services.

The universities will deliver the satellites to AFRL inApril 2001 for integration onto the deployment platform.The integrated payload will then be delivered to thelaunch vehicle integrator in June 2001 for launch on orafter November 2001. Further information on theprogram and presentations from the kick-off meeting canbe seen at the website, http://www.nanosat.usu.edu/.

The universities selected for the program are: ArizonaState University, University of Colorado at Boulder, andNew Mexico State University (Three Corner Sat);Stanford University and Santa Clara University(EMERALD); Utah State University, VirginiaPolytechnic Institute and State University, University ofWashington (ION-F); Boston University (ConstellationPathfinder); and Carnegie Mellon University (SolarBlade Nanosat). Descriptions of their satellite programsfollow.

Three Corner Sat Constellation (3ÙSat)Arizona State University: Brian Underhill, Assi

Friedman, Helen ReedUniversity of Colorado Boulder: Elaine Hansen, Dan

Rodier, Anthony ColapreteNew Mexico State University: Stephen Horan

Overview

This project is a joint effort among Arizona StateUniversity (ASU), University of Colorado at Boulder(CU), and New Mexico State University (NMSU). Aptlynamed Three Corner Sat (3ÙSat), our proposedconstellation of three identical nanosatellites willdemonstrate stereo imaging, formation flying/cellular-phone communications, and innovative command anddata handling. In addition, each University in our 3ÙSatconstellation has the opportunity to fly an individualunique payload should it desire. Because of our team’sheritage in space flight (CU’s DATA-CHASER payloadvia August ’97 Space Shuttle), conventional satellitedesign (CU’s Citizen Explorer via Delta, December ‘99),and nanosatellite design (ASU’s 4.5kg ASUSat1 via OSPMinotaur, September ’99), our constellation will beready for launch in late 2001.

1. Mission Goals

Stereo Imaging. The primary science objective of the3ÙSat constellation is to stereo image small (< 100meter), highly dynamic (< 1 minute) scenes includingdeep convective towers, atmospheric waves, andsand/dust storms. These stereo images will enable thecomputation of range to within 100 meters givingaccurate data regarding the shape, thickness and heightof the observed phenomena.

Stereo imaging from space has several advantages overconventional imaging, the most obvious being the abilityto derive range data. This range data can be substantiallymore accurate than range data acquired by other moreusual means and also can cover a much greater area.Stereo imaging involves correspondence matchingbetween an image pair and calculation of the resultingdisparity. From the disparity, triangulation can be used todetermine range data, and three-dimensional images anddepth maps can be created. Accurate depth maps withrange resolutions of about 100 meters enable the study ofrelatively small-scale, short-lived atmospheric eventssuch as cumulus-cloud towers.

Zones of deep convection, in areas such as the Midwest,frequently create large cumulus towers that extend from

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the middle troposphere into the lower stratosphere. Thesezones of convection are frequently impassable to airtraffic due to the highly unstable air. Radar, while able towarn aircraft of large convection cells, is unable to giveaccurate data as to their extent in altitude. Thus, airtraffic is frequently diverted hundreds of miles regardlessof the altitude extent of the convection cells. With betterestimates of cloud heights, some air traffic may be ableto traverse these convective boundaries by flying overareas of shallow convection.

Cloud heights are critical to our understanding of theEarth’s climate and our ability to better model it.Because of their dynamic nature, both spatially andtemporally, incorporating clouds and their effects intoGlobal Circulation Models has been difficult. One keypiece of data that is missing is the height and thickness ofclouds at a global level. Using stereo imaging, we willmeasure the heights of clouds with a precision of lessthan 100 meters and make a statistical study of their type,height, and thickness.

In the last decade, studies have indicated that just asimportant as clouds, other aerosols, including mineralaerosols such as dust or sand, play an important role inEarth’s climate system. Recent experiments have beenundertaken or planned to understand the composition,structure, and distribution of mineral aerosols on both alocal and global scale. Stereo imaging allows thestatistical study of aerosol cloud structures, such as sandstorms, and can provide information on the relationshipbetween uplift efficiencies, boundary-layer thickness,and particle sizes with local environments.

As the time between typical satellite images can be long,highly dynamic objects such as clouds and dust stormsare currently stereo imaged in a way that makes therange data inaccurate. Stereo images from the GOES8and GOES9 weather satellites have proven theeffectiveness of using two satellites to view the samescene, however, these satellites can only view togetherfor a few hours a day, and, since they have relatively lowimage resolution, the range data is poor. For highlydynamic objects, several satellites with relatively goodresolution need to image the same location at the sametime. By using a formation of satellites, stereo images ofsmall, highly dynamic objects can be made, and fromthese stereo images, accurate range data may becalculated.

Formation Flying / Communications. To accomplishthe science objectives, a “virtual formation” is proposedand will be demonstrated as part of our program. Thevirtual formation is a cooperative effort betweensatellites operating as a network where targeting and dataacquisition are accomplished and results transmitted tothe ground segment and to the other satellites viacommunications links without the need for strict physicalproximity of the satellites. In this mode, thecommunications links carry the command and controldata necessary to accomplish the mission regardless ofthe physical location of the satellites. For the mission to

be accomplished, the locations of the satellites will needto be “in range” and mutually known in order for each tosupport its portion of the mission, but physical proximityis not a requirement for the formation network.

For stereo imaging, a nominal spacing of tens ofkilometers between the satellites is required. With acontrolled deployment to achieve this initial spacing, thesatellites will remain within range for the suggested four-month lifetime of the mission. Therefore propulsivecapability is not needed.

The design of the mission utilizes a commercialcommunications network in Low Earth Orbit (LEO)which supplies the communications links as shown inFigure 1. This will allow each satellite to be contactedvia the LEO network regardless of the position of thesatellite relative to the ground station – with predictablevisibility outages. Because each satellite in the networkwill be visible to the LEO communications constellation,there will be the ability for satellites to perform theirmission coordination without the need for visibility fromthe ground station or with each other. The LEOcommunications network knits together the virtualformation.

LEO satellites utilizing cellular telephone constellationsis a new concept but one in which there is considerableinterest in the government and private-sector spacecommunities. This natural extension to the use ofground-based systems will be explored not only todemonstrate the utility of this mode of communicationsbut also to act as an experiment to characterize theconstellation itself and the limits on the operations. Atechnology goal of 3ÙSat is to perform the first steps inthis characterization.

Command and Data Handling (C&DH) System. TheC&DH for the 3ÙSat constellation is designed as adistributed and simple system. As part of this distributedarrangement, each satellite uses a Satellite ProcessorBoard that serves as its local controller, data interface,on-board memory, and processor. The three-satelliteconstellation can be controlled and managed by aprocessor on any of the three satellites via thecommunication links. The Satellite Processor can beresponsible for supervising the operation of the threespacecraft and managing their resources. Thissupervision can be automatically accomplished withinthe constellation by the selected satellite processor whichcan initialize and distribute commands and which canmonitor and react to science and engineering data fromthe three spacecraft.

University Specific Experiment – ASUMicropropulsion System. Micropropulsion systems canoffer a wide variety of mission options, all relevant toformation flying: attitude control, orbital drag make-up,altitude raising, plane changes, and de-orbit. For itsUniversity-specific experiment, ASU is collaboratingwith AFRL and industry to design and fly amicropropulsion system. The objective of ASU’s

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research is to take a systems point of view and develop asafe and simple micropropulsion system fornanosatellites. In particular, the ASU satellite willdemonstrate orbit raising and de-orbiting once the 3ÙSatvirtual-formation/stereo-imaging mission is completed.

2. Mission Implementation

Mission Design. The 3ÙSat constellation will consist ofthree satellites flying in a linear follow-formation withrelatively constant separation from each other. Theseparation distance selected is based on altitude andcamera field of view (FOV), with final determinationbased on the chosen launch vehicle. The mountingconfiguration within the launch vehicle will depend uponthe launch vehicle and other satellites selected. Thesatellite will use gravity-gradient (GG) forces forstabilization with +/- 5 degrees pointing accuracy.

Spacecraft. All satellites will be identical, except for astandard payload envelope where each university willhave the option to fly its own unique experiment after theprimary science objective has been met. The spacecraftstructure will be low-cost and reliable. The exteriorenvelope of the structure is a six-sided disk structureconsisting of tubular supports and machined end caps tohold the bulk of the loading (Figure 2). The design willfeature a number of modular, removable trays, allowingfor on-the-spot modifications without extra machining orirreversible processes. The design incorporates acommon electrical bus that is easily accessible anddurable. The material selection and componentplacement will eliminate the need for any extra shieldingfrom EMI or radiation in space. The solar array panelswill be mounted on thin aluminum sheets that mount tothe exterior of the frame. All components will mount toaluminum honeycomb plates, which fasten to the mainframe via slide-in interface brackets, and/or standardsocket head cap screws. The batteries will be stored inthe middle of the structure to avoid an unbalancedinertial configuration. These cells will be housed in aneccofoam/aluminum structure attached in a manner tostiffen the component panels from harsh vibrationenvironments.

Project Schedule

Milestone DateProject Start January 1999System Requirements Review May 1999Technical Interchange Meetings (every 2 months)Preliminary Design Review August 1999Critical Design Review March 2000Build March-August 2000Integration and Test Fall 2000Environmental Test November 2000Qualification Readiness Review December 2000Flight Readiness Review February 2001Launch-Vehicle Integration April 2001Launch late 2001

3. Sponsors

AFOSR/DARPA, NASA Space Grant Program,Lockheed Martin Tactical Defense Systems, Honeywell,Lockheed Martin Astronautics, Motorola, KinetX,AFRL, Cogitec, Space Quest, Microchip, BallAerospace, Spectro Lab, and JPL.

C O N M A Z

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L E O C o m m u n i c a t i o n C o n s t e l l a t i o n

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C o n c u r r e n tCons te l la t ion View

C o n t r o l C e n t e r s a n d D a t a A r c h i v e s

~ 3 0 k m

Figure 1. 3^Sat Constellation Overview

Primary exterior components shown:1. Phone antenna2. Star mapper (15O FOV)3. Parallel gravity gradient booms with tip masses4. Integrated battery pack/release mechanism5. GPS patch antenna6. GaAs body mounted solar array (18% efficient)7. Hard mounting points / lateral movement restraint8. Four CMOS cameras (FOV 15O single/54O composite)

Primary interior components not shown:1. Boom deployment mechanism2. C&DH electronics3. Power control board4. Cell phone5. GPS receiver6. Paraffin actuated pin-pullers7. Structural supports, tubes and panels8. Camera boards with micro controllers

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Figure 2. 3^Sat w/Launch Configuration

EMERALD: A LOW-COST SPACECRAFTMISSION FOR VALIDATING

FORMATION FLYING TECHNOLOGIES

Bob Twiggs - Stanford Space Systems DevelopmentLaboratory

Chris Kitts - Santa Clara Remote Extreme EnvironmentMechanisms Laboratory

Jon How - Stanford Aerospace Robotics LaboratoryStudent Managers: Bryan Palmintier and Freddy

Pranajaya

Spacecraft formation flying is an evolving technologywith vast scientific, military, and commercial potentialsthat range from enhanced mission performance to radicalreductions in operations cost. As part of the TechSat 21University Nanosatellite Program, Stanford Universityand Santa Clara University are jointly developingEMERALD, a low cost, two-satellite mission forvalidating formation flying technologies.

Stanford’s SSDL (Space Systems Development Lab) andSanta Clara University’s SCREEM (Santa Clara RemoteExtreme Environment Mechanisms Laboratory) willwork as a unified team to develop, construct, test andeventually operate the EMERALD spacecraft. Theformation flying experiments will be coordinatedthrough Stanford’s ARL (Aerospace Robotics Lab)

The EMERALD Mission is divided into three distinctstages that progress from a simple single satellite to twofree flying satellites in a coarse formation. Using abuilding block experimental strategy, the researchpayloads first will be characterized in isolation. Then,they will be coordinated and combined to permit simpledemonstrations of fundamental formation flying controlfunctions such as relative position determination andposition control.

• At release, the two spacecraft will be stackedtogether and will travel as a single object. This willallow initial checkout, calibration, and some limitedexperimentation.

• During the second stage of operation, the satelliteswill separate and a simple tether or flexible boomwill uncoil, linking the two vehicles. This tetheredstage will allow full formation flyingexperimentation including on-orbit relative positiondetermination and simple closed loop relativeposition control using the drag panels andmicrothrusters.

• During the final stage of operation, the tether will becut in order to permit true two-body formation flyingfor a limited period of time. The tether will have asimple sub-satellite at its midpoint. Upon groundcommand, the two halves of the sub-satellite willseparate. Each satellite will retain half of the tetherand half of the sub-satellite, providing very roughgravity gradient stabilization.

Figure 3. EMERALD Mission Sequence(Joined, Tethered, Formation Flying)

The two EMERALD spacecraft will demonstrateseveral critical technologies for future formationflying missions:

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• GPS-based positioning For onboard orbitdetermination and relative navigation, a Stanford-modified Mitel 12-channel GPS receiver will beflown on each spacecraft. This receiver can computerelative position to approximately 2-5 meter levelaccuracy in real-time using differential GPStechniques. They will be modified-for-spaceversions of the receivers currently in use by ARL’sother formation flying experiments.

• Inter-satellite communication EMERALD plans todevelop a simple inter-satellite communication linkfrom the commercially available19.2 kbs wirelessradio modems currently used by ARL for groundbased formation flying systems. This will providethe real-time communication link necessary fordifferential GPS measurements.

• Advanced colloid microthrusters To enable smallscale position control, these thrusters can supplyvectored thrust on the order of 0.11 milli-Newtons,and have a specific impulse of approximately 1000seconds. These components are currently indevelopment by Stanford’s Plasma DynamicsLaboratory (PDL).

• Simple, passive position control devices

- A simple tether or flexible boom will maintainthe satellites within a given distance, on theorder of tens of meters. This tether may be cutlater in the mission in order to demonstrateadvanced formation flying capabilities.

- Deployable panels on both spacecraft will allowsimple, low performance drag control. Duringthe tethered mission phase, these panels can beused to maintain tether tension as well as toattempt closer positioning.

In addition to these mission critical payloads,EMERALD will support a couple of auxiliary payloads:

• VLF receivers will record radio waves emitted byhigh altitude atmospheric phenomena associated

with lightning. The experiment will help understandsome of the fundamental happenings in the EarthIonosphere. The distributed sensing enabled byformation flying capabilities, will enhance thescientific data and help validate the benefits ofsatellite formation flying. The receivers will bebased on an existing design developed at STAR Lab(Stanford Space Telecommunications andRadioscience Laboratory) The mission name,EMERALD (Electromagnetic Radiation AndLightning Detection), refers to this experiment.

• MERIT, the MicroElectronics Radiation In-flightTestbed, will characterize the performance ofadvanced microprocessors, MEMS technologies,and other electronic components in the spaceenvironment. This payload is being developed aspart of a separate SSDL research program inconjunction with Boeing, the Naval ResearchLaboratory (NRL), The Aerospace Corporation,Honeywell, UTMC, and the Laurence BerkeleyLabs.

These experiments will be built into a 12-inch tall, 18-inch diameter hexagonal bus adapted from heritagedesigns at SSDL. The structure employs a modular,stackable tray structure made of aluminum honeycomb.Inside, command and data handling is provided by aMotorola 68332-based processor built into a studentdeveloped radiation tolerant architecture. Groundcommunication will be handled by a 9.6 kbs packetcommunications system. The heritage power systemconsists of body mounted Silicon cell solar panels and asingle Ni-Cad battery. This system will provide 7 Wattsof average power to components via a 5V regulated bus.Passive thermal control will be achieved through the useof insulation and thermal coatings.

(a) Assembled View (b) Exploded View.

Figure 5: Heritage Bus Configuration

One of the most important upgrades to the heritagearchitecture is the shift to a serial command and data bus.As shown below, EMERALD will use I2C to connect all

Figure 4. Mitel-1GPSReceiver (Top left), ARLRadio Modem, ColloidMicrothruster

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of the subsystems and experiments. This modular designsimplifies integration and allows stand-alone testing ofcomponents. It also allows the project to maintain aschedule driven design: If a payload fails to show enoughprogress at a quarterly design review, it can be removedfrom the mission, with minimal effect on the rest of thesatellite. For future projects, this modular bus structurewill allow subsystems and experiments to be selectivelyswapped-out, upgraded, and interchanged, similar toreconfiguring a PC.

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Figure 6: EMERALD System Block Diagram

Ionospheric Observation NanosatelliteFormation (ION-F)

Utah State University: Dr. Frank Redd (PI),Dr. Charles Swenson, Dr. Rees Fullmer

University of Washington: Dr. Mark Campbell (PI),Dr. Adam Bracken

Virginia Polytechnic Institute & State University:

Dr. Chris Hall (PI), Dr. Nathaniel Davis, Dr. JaimeDeLaRee, Dr. Wayne Scales, Dr. Warren Stutzman

Program OverviewUtah State University, University of Washington, andVirginia Polytechnic Institute are designing anddeveloping a system of three 10-kg spacecraft toinvestigate satellite coordination and managementtechnologies and distributed ionospheric measurements.The three will coordinate on satellite design, formationflying and management mission development, andscience instruments and mission. Advanced hardwarefor distributed space system to be demonstrated includesm-pulsed plasma thrusters, gimbaling magnetic attitudecontrol, and an advanced tether system. In addition, anInternet based operations center will be designed toenable each university to control its satellite from its ownremote location. ION-F will focus on mission objectivesthat would benefit TechSat 21 and future Air Force andNASA missions. Coordination between the differentuniversity teams will be by conferencing, Internet, andtelephone. In addition, industrial (SDL, Primex,Honeywell) support has been identified, including

funding for students, hardware, and satellite designsupport.

Objectives and Research Issues

The program objectives (PO’s) for the ION-F follow:

PO1. Basic research mission of investigating globalionospheric effects which affect the performance ofspace based radars, and other distributed satellitemeasurements.

PO2. Formation flying and local communication in aconstellation, including upgrade from a 3 nanosatelliteconstellation to 4 nanosatellites.

PO3. Baseline new technologies including micro-thrusters, magnetic gimbaled attitude control, advancedtether system, and an Internet based operations center.

PO4. Bring a unique, hand-on space experience tograduate students, undergraduate and graduate spacedesign students, and undergraduate engineering studentsin independent research projects.

Formation FlyingBecause formation flying is a primary mission objectivefor ION-F, and because at least one satellite will havepropulsion capability, the UW-VT-USU team is veryinterested in working with the NASA Goddard formationflying program. Communications cross-links are the toppriority, and the relative navigation chip and software areimportant as well. Because of the tight constraints andhigh capability of the ION-F nanosatellites (includingpropulsion), there is little power or mass budget foradditional instruments in the preliminary designs. TheUW-VT-USU team intends to fly their control andformation algorithms but are investigating additionalcollaboration with NASA Goddard and their partners.

Figure 7: Ionospheric Observation Nanosatellites

It is expected that each of the three satellites will havecross-communication links and relative navigationcapabilities. In addition, UW will have propulsion

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capability and USU will attempt to use differential drag.VT is undecided at this time, but is examining AFRLmicro-propulsion concepts. With this in mind, thefollowing activities were outlined as part of theformation flying mission of ION-F:

• After deployment of the three linked satellites, acheckout will occur of the subsystems, including GPScalibration, attitude determination, and possiblycommunications.

• After initial checkout and relative calibration, thesatellites will separate. The satellites will deploy into aclose leader-follower formation, and individualperformance characterization and disturbancequantification will be performed.

• Using UW/Primex Micro-Pulsed Plasma Thrusters(µPPT’s) and differential drag capability of USU, two-satellite formationkeeping will be accomplished relativeto the VT satellite.

• More complex two-satellite formations will beexamined such as side-by-side (same altitude butdifferent inclination) and same ground track (NASAGoddard’s “ideal” formation). The operations willinclude maneuvering into a new formation, andsubsequent formationkeeping.

• Complex three-satellite formations will be attempted.Two examples include 1) maneuvering three satellites ina leader-follower formation to three satellites with thesame ground track and 2) a rotating formation about anequidistant point.

• Formationkeeping using both position and attitude.

• Additional NASA Goddard collaboration and controlalgorithms can be accommodated.

Distributed Ionospheric DataThe objective is the understanding of ionospheric densitystructures that can impose large amplitude and phasefluctuations on radio waves passing through theionosphere. The constellation provides a uniqueopportunity to answer questions about ionosphericdisturbances that can not be addressed any other way.A single satellite can only provided very limitedinformation on the dimensions and evolutionary timescales of the ionospheric disturbances it flies throughbecause a full orbit (90 minutes) must occur between thenext observation. In general the situation is even worsethan this because only truly zero inclination equatorialsatellites have a good possibility of measuring the sameregion twice due to the co-rotation of the ionosphere withthe Earth. This science investigation contributes to theTechSat 21 basic research mission of investigating globalionospheric effects which affect the performance ofspace based radars. It also addresses broader Air Forceinterests in ionospheric effects on navigation andcommunication links.

We propose using the nanosat constellation to make thefirst global multi-satellite electron density measurementsin the ionosphere. We also propose to make the first

global multi-baseline RF-scintillation measurements ofthe ionosphere. The scintillation of GPS signals usingreceivers on each spacecraft will provide informationabout the scale sizes of disturbances between thenanosatellite constellation and the GPS transmitter.

The Earth's ionosphere frequently shows densitydisturbances and fluctuations over a very large range ofscale sizes (from hundreds of kilometers to centimeters)at all latitudes, longitudes and nearly all altitudes.Ionospheric plasma density irregularities are associatedwith two-dimensional turbulent processes driven by bothlow latitude tidal neutral winds and high latitude currentsystems. We propose to determine the global distributionof plasma structure in both the quite and disturbedionosphere. The proposed measurements would provideessential information for the understanding ofionospheric effects on communications, navigation, andGPS systems and for the development and validation ofrealistic predictive global ionospheric models. Theseobservations are related to the priority measurements setup the by the National Space Weather Program.

One of two possible instruments will be deployed oneach of the satellites in the constellation, either a plasmafrequency probe from which an absolute electron densitycan be determined or an electron saturation current probewhich can measure relative electron density variations.The electron saturation current probe is based on the DCresponse of a plasma to an applied potential on a probe.These types of probes were pioneered in the early 1920sby I. Langmuir and consequently are also known asLangmuir probes. The plasma frequency probe techniqueis based on the AC response of the probe. Early work inthis technique can be traced back to diagnostics on V2rockets shortly after World War II. These types of probeshave been called impedance probes, RF-probes, andcapacitance probes. The instrument to be selected will bebased on the resources available on the nanosat but ingeneral these instruments will be small, low-power, andhave low impact on their host satellites.

The scintillation measurements will be extracted fromthe GPS receivers that are part of the orbit determinationsystem on the nanosats. The 1575 MHz signal from theGPS satellites originate at 20,000 km over the Earth andmust travel through the ionosphere, line of site, to thelocation of the nanosats at 360 km altitude. The signalwill encounter regions of disturbed ionospheric plasmawhich will slightly increase or decrease the signalstrength at the receivers. The size of these disturbedregions can be estimated by comparing signals measuredover closely related propagation paths, such as betweentwo nanosats.

Satellite Design

The basic satellite design is configured for released froma Shuttle Hitch-hiker canister and will also accommodateAFRL’s proposed deployment from the shuttle SHELS.The structure is hexagonal with an approximate 18 in.width and a 5 in. height. This configuration can be

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modified as needed to meet the specific launchenvironment conditions. The structure will consist of twoaluminum hexagonal plates and six end covers. One ofthese plates will serve as the mounting base plate for thespacecraft. The power subsystem will be a typical solarcell/battery design using high grade commercial industryNiCd battery technology. The entire spacecraft will becovered with solar cells, with the exception of the areasdesignated for the sensors. The power regulationelectronics will leverage high quality commercial NiCdbatteries which have been developed commercially forportable laptop computers.

Advanced HardwareIn addition to the science mission and formation flyingexperiments, at least one new technology will also bebaselined on each of the satellites in ION-F. Fourpossible technologies, and their researchers, aredescribed subsequently. The process will include keepingthe collaborators up to date with the nanosatellite design.During the critical design phase of the program, adecision will be made as to the maturity of eachtechnology for the nanosatellites, and whether it can beflown. Because of complexity and reliability issues, thesetechnologies must have a high probably of success.Currently, the new technology options include:microthrusters, magnetic control, tethered deployment,satellite cross-links, and Internet operations.

Two versions of microthrusters are currently indevelopment. Primex Aerospace Company is workingwith the University of Washington to scale down thepower requirements of their Pulsed Plasma Thrusters.The UW nanosatellite will fly a propulsion system, andwill be either the m-PPT’s, or a cold gas system. Primex,Honeywell and AFRL are working separately on MEMSbased thrusters such as micro-hydrazine. These will beflown on either the USU or VT nanosatellite if thematurity of the technology will allow it. The smallmodular nozzles would allow many options as tomicrothruster size. Although development time will mostlikely require more than two years, the potential fornanosatellites is very high.

Figure 8: Full-sized Pulsed Plasma Thrustersfrom Primex Aerospace Company

Magnetic Control: One objective is to develop 3-axisattitude control given the very limited power and weightavailability on a nanosatellite. We will meet thischallenge with an all magnetic torquer system wherepermanent magnets on stepper motors are used instead oftraditional torquer coils. The attitude determination willbe achieved by a combination of Earth horizon and sunsensors, giving three-axis control to approximately twoto three degrees.

Tethered system: Adding a low-mass tether for gravitygradient stability will provide simple attitude control forthe VT satellite. Nominally the tether will be a simplenon-conducting ribbon. The spacecraft will also includea small digital camera for imaging the tether duringdeployment and during the eclipse exit period of theorbit. This will provide useful data on the flexibledynamics of tether systems.

Satellite Cross-links: Collaborating with NASAGoddard and building on the past experiences of USUand the SDL, the ION-F team will demonstrate satellitecross-links, exchanging relative GPS information andpossibly attitude information.

Ground Operations via Internet: The USU groundstation will be used for ION-F, and another groundstation, most likely at VT, will be developed. In order tosuccessfully coordinate the three satellites, an Internetbased ground station will be developed. This will allowcoordination of experiments by the different universitiesas well as individual satellite control and datadissemination. VT is also investigating use of theGlobalStar LEO constellation for satellite telemetry,tracking and commanding of the satellite usingcommercial communications technology.

CONSTELLATION PATHFINDER: A UNIVERSITY NANOSATELLITE

BOSTON UNIVERSITY AND DRAPER LABDR. HARLAN SPENCE

MISSION OBJECTIVE

The Constellation Pathfinder program demonstratesthe feasibility of fabricating and launching one tothree, <1 kg satellites that are capable of collectingand returning quality scientific and engineering datafor one to four or more months . The particular satellitewe propose to use is based on one developed over thepast two years through a NASA-supported study calledthe Magnetospheric Mapping Mission (MMM) at BostonUniversity. That study objective has been to assess thefeasibility of placing hundreds of satellites equipped withmagnetometers, into orbits extending into the tail of themagnetosphere, thereby obtaining a much more detailedthree-dimensional picture of dynamic phenomena ingeospace than has been possible previously. TheConstellation Pathfinder proposal will take the firstpathfinding step toward such an ultimateimplementation. We propose a simplification of ourcurrent conceptual design in that the launch mechanism

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RF Antenna

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is provided by the Shuttle Hitchhiker, the magnetometerwill be measuring larger (and therefore easier tomeasure) fields in the Earth’s ionosphere, the loweraltitude reduces RF communication requirements as doesrelaxation of the required data transmission rate, and thenatural radiation environment will be much lower. Thehardware demonstration of building and flying such asatellite, or small suite of satellites, will provide aproof of principle that will be helpful in manyscientific and strategic applications where a fleet ofcoordinated small satellites is required.Demonstration of satellite-to-satellite communicationmay also be possible.

SATELLITE AND PAYLOAD DESIGN

The nanosatellite configuration and functional blockdiagrams are shown schematically below. The outercylindrical surface consists of power-providing GaAssolar cells. The RF antenna is along the axis of thecylinder. Satellite spin will maintain the orientation ofboth the antenna and the solar cells within 10 or 20degrees of the ecliptic plane assuming that we can selectthe shuttle orientation at the time of release to be in thisrange. The satellite electronics including batteries will beconcentrated in the center. The magnetometer location ischosen to minimize contamination by spacecraftmagnetic fields and will be located within the cylindricalvolume. The sun sensor looks radially outward and willbe used to determine the phase of rotation. Additionallyin conjunction with the magnetometer it will determinethe direction of the spin axis.

Figure 9: Pathfinder Satellite and Functional Layout

As compared to typical satellite designs this mission isparticularly stringent in terms of requiring low mass andlow power. In view of the large numbers of satellites

eventually involved, the design must addressmanufacturability – simplicity of fabrication, assemblyand calibration. On the other hand, the large number ofsatellites also reduces the reliability requirements.Failure of a few satellites simply reduces the number ofdata points but it does not lead to mission failure.

SOLAR BLADE HELIOGYRO NANOSATELLITEDr. Richard Blomquist and Dr. William Whittaker

Carnegie Mellon University

Solar sail concepts have existed for decades, but theirimplementation has been elusive, and none have flown.The primary difficulty has been the need for greatsurface area relative to mass. Traditional spacecraftdesigns with hundreds of kilograms of mass led tokilometers of blade dimensions, which were impossibleto rationalize, build, and fly. Nanosat technologydrastically reduces mass and makes heliogyro designeminently practical and flyable.

Carnegie Mellon proposes to develop and fly the firstsolar sail, a spacecraft which utilizes solar radiationpressure as its only means of propulsion and attitudecontrol. The solar pressure will enable changes toaltitude, attitude precession, spin rate and orbital position.

The Solar Blade Heliogyro Nanosatellite has theappearance of a Dutch windmill and employs controlakin to a helicopter. Four solar reflecting blades mountradially from a central spacecraft bus and actuate alongtheir radial axis. The satellite uses collective and cyclicpitch of these solar blades relative to the sun’s rays tocontrol its attitude and thrust. The spacecraft weighs lessthan 5 kilograms, and, when stowed, is a packageapproximately the size of a fire extinguisher.

The satellite will demonstrate attitude precession, spinrate management, and orbital adjustments, after which itwill spiral out past the orbit of the moon. For the SolarBlade Nanosat, plane change maneuvers will be most

Figure 10. Solar Blade Heliogyro Nanosatellite

10

efficient when the sun is furthest out of the orbit plane.This increases the magnitude of the orbit-normalcomponent of force that can be used for the plane changemaneuver. Plane change maneuvers can also beconducted if the sun lies in the orbit plane by orientingthe solar blades at an angle relative to the orbit plane,optimally 45°. Unlike eccentricity changes, which canbe implemented throughout the orbit using a single solarblade orientation, plane change maneuvers must changepolarity on opposite ends of the axis of plane rotation.This is not possible unless the sun is in the plane of theorbit (the solar blades cannot produce a positive orbitnormal force if the sun is above the orbit plane).Therefore, in most situations, plane change maneuverswill be conducted over an orbital arc on one side of theorbit near the axis of desired orbit rotation. In addition toattitude and orbital maneuvering, the ultra-lightspacecraft will communicate with the Earth, uplinkingcommands and relaying orbital and attitude informationto ground stations.

The Solar Blade Nanosatellite consists of a corecontaining the computer, communications system, andattitude determination hardware. Four bending strutsemanate from the core and solar cells cover the topsurface. The blades attach to the struts throughindividual actuators.

Each blade of the Solar Blade is a 20 meter long by 1meter wide aluminized Kapton sheet 8 microns thick.Edge reinforcing Kevlar and battens of 80 micron-thickKapton provide added stiffness and resistance to tears.Small brushless motors rotate the blades.

Solar cells embedded on the flanges of the C-beam frameprovide up to 28 Watts of power. The spacecraftcomputer, communication system, and station-keepingsensors at the center of the square connect to the framethrough thin lenticular beams. Wiring between the solarcells and the other subsystems consists of thin-filmflexible printed circuits.

Deployment Structure DesignAndrew Peffer and Jeff Ganley, AFRL/VSD

The AFRL Space Vehicles Directorate will design,fabricate, and test a deployment structure to eject fromthe launch vehicle and deploy the 10 universitynanosatellites. Two preliminary designs have beenformulated to accommodate launch from an expendablelaunch vehicle as a secondary payload or from theShuttle Hitchhiker Experiment Launch System (SHELS).The SHELS is a side bay ejection system underdevelopment that accommodates significantly greaterweight and volume (180kg and 54” x 42.5” x 24”) thanthe typical hitchhiker canisters.

The primary objective for the deployment structure is to1) provide an interface to the shuttle SHELS system and2) eject the nanosatellites following separation from the

shuttle. It will not provide orientation, communication,or spin up capability, and the nanosat separations areexecuted via timer using battery power.

This project poses a unique problem with regard to thedesign of a deployment structure. Since each of theuniversities has unique configurations within volume andweight constraints, this increases design complexity forload transfer to the supporting structure and forarrangement within the launch vehicle. Preliminarydesign concepts for launching on 2 SHELS or on thePegasus or OSP Minotaur follow.

Figure 11. Shuttle Hitchhiker ExperimentLaunch System (SHELS)

42.5”

BU24”

USU3^Sat

35”

ASU

UCB

USU

UW

VTBU

42.5”

NMSU

Figure 12. SHELS Payload A Configuration

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Conclusion

The 10 university nanosatellites provide a broad range oftechnology demonstrations in the areas of miniaturespacecraft subsystem components and formation flying.

There are also numerous science measurements andexperiments in such areas as GPS scintillation, solarwind, magnetic fields, and upper atmosphere ion density.

This program has the potential to provide significantpayoff for very modest funding by DoD and NASAgiven the broad university resources being applied andsupport by industry partners. If these flightdemonstrations are successful, it is very likelygovernment sponsorship can be secured for follow-onlaunches of nanosatellites built by universities and otheragencies.

Figure 12. SHELS Payload B Configuration(Includes Stanford ORION Satellite funded by

NASA Goddard)

3 CornerSat.

3@18”Dx 10”H

EMERALD

2@18”Dx 16”H

USU,VT,UW

3@18”Dx 10”H

SolarBlade

12”x12”x36”

Pathfinder3@ 7” Dx 2.5”H

46”

36” height

Figure 13. Pegasus/OSP MinotaurLaunch Vehicle Configuration

Sep Sep

Stanford

SantaClara

37”

42.5”

EMERALD

42.5”

StanfordORION

StanfordORION