Development of Sample Verification System for Sample ... - NASA

Development of Sample Verification System for Sample Return Missions Risaku Toda*, Colin McKinney, Shannon P. Jackson, Mohammad Mojarradi, Ashitey Trebi-Ollennu and Harish Manohara† Jet Propulsion Laboratory, California Institute of Technology 4800 Oak Grove Drive Pasadena, CA 91109 818-393-5483 * [email protected], †[email protected] Abstract—This paper describes the development of a proofof-concept sample verification system (SVS) for in-situ mass measurement of planetary rock and soil sample in future robotic sample return missions. Our proof-of-concept SVS device contains a 10 cm diameter pressure sensitive elastic membrane placed at the bottom of a sample canister. The membrane deforms under the weight of accumulating planetary sample. The membrane is positioned in proximity to an opposing substrate with a narrow gap. The deformation of the membrane makes the gap to be narrower, resulting in increased capacitance between the two nearly parallel plates. Capacitance readout circuitry on a nearby printed circuit board (PCB) transmits data via a low-voltage differential signaling (LVDS) interface. The fabricated SVS proof-of-concept device has successfully demonstrated approximately 1pF/gram capacitance change.1, 2

return phase of the mission is initiated. For some robotic sample return missions sample acquisition verification must be done autonomously without ground in the loop of operations. There is a technology gap for sample acquisition verification systems for robotic sample return missions. NASA's Genesis and Stardust robotic sample return missions successfully returned samples to Earth without an in-situ sample acquisition verification system (SVS) on board the spacecraft [1,2]. Positive confirmation of successful sample acquisition and transfer was done after the return of the sample capsule to Earth. These two missions are an exception because of the types of sample they acquired, no direct interaction with the target body was required and sample acquisition time was in order of several minutes. JAXA's Hayabusa mission also did not have an insitu sample acquisition verification system; as a result, positive confirmation of successful sample acquisition and transfer could only be done after the return of the sample capsule to Earth [3]. The approach of providing positive confirmation of successful sample acquisition and transfer for robotic sample return missions after the return of the sample capsule to Earth is less than ideal since the ultimate goal of a sample return mission is to return an assured sample quantity (threshold science) to Earth.

TABLE OF CONTENTS 1. INTRODUCTION .................................................................1 2. BACKGROUND ..................................................................1 3. SENSOR DESIGN ................................................................2 4. PROOF-OF-PRINCIPLE TEST ............................................4 5. CAPACITANCE READOUT CIRCUIT ..................................5 6. SUMMARY .........................................................................5 ACKNOWLEDGEMENT ..........................................................6 REFERENCES ........................................................................6 BIOGRAPHY ..........................................................................7

This point argues for additional technology development for future sample acquisition verification systems for potential robotic sample return missions. In this paper, we present a novel in-situ SVS that is designed for integrated sampling systems that could survive and operate in challenging environments (extremes in temperature, pressure, gravity, vibration and thermal cycling) for real-time in-situ sample acquisition verification. The in-situ SVS would enable the unmanned spacecraft system to re-attempt the sample acquisition procedures until the capture of desired sample quantity is positively confirmed, thereby maximizing the prospect for scientific reward.

1. INTRODUCTION In previous human lunar sample return missions (Apollo 11, 12, 14, 15, 16 and 17), astronauts were able to ascertain the quantity of lunar samples before returning to Earth. Robust in-situ sample acquisition verification (assured sample quantity (mass or volume)) systems would be critical to the next generation NASA robotic sample return missions. For sample return missions the Sample Transfer Chain (STC) would be responsible for the acquisition, verification, containment and transfer of the sample from the planetary body to the surface of the Earth. A key mission success criterion for robotic sample return missions would be that an assured sample quantity has been collected before the Earth1 2

2. BACKGROUND The primary objective of a sample return mission is retrieving pristine planetary sample without contamination. Spacecraft systems including the SVS must be designed so

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atmosppheric entry, deescent and lannding phases uupon return to Earthh.

thhat it will not contaminate th he planetary saample. Even veery ssmall contamin nation may im mpact accuraccy of elemen ntal aanalysis such as a isotopic age determinatio on. For examp ple, isotope compo osition analysis of Apollo 11 1 lunar samp ple rrequired 0.005 5%-level precision [4]. Trad ditionally, a few fe m materials such as stainless steel s SUS 304 4 and Aluminu um 66061 alloy hav ve been used for planetary sample handliing aapplications. Therefore, T we use u these mateerials exclusiveely ffor areas that may m become con ntact with the sample. P Planetary/Moon n environmen nt is harsh for spacecrraft ccomponents. Temperature T ch hange may excceed 100 Kelv vin dduring a day mainly m due to change of solaar radiation. The T S SVS sensor mu ust survive succh a large tem mperature chang ge. C Careful choice of materials and a design is necessary to bu uild a robust SVS sensor s that is stable s against large temperatu ure cchanges. Addittionally, the SV VS sensor sho ould be design ned sso that planetaary atmospheree and/or vacu uum environmeent w will not impactt sensor perform mance. T There are several potential approaches to detect samp ple ppresence in ro obotic spacecraafts. For exaample, (1) Maass m measurement by b spring-mass system and transducers, (2) ( O Optical metho ods such as photo imagiing, (3) Inerrtia m measurement by b Newton's seecond law of motion. m The SV VS ssensor develop pment describeed here is inteended for planets aand moons wh here significantt gravity existts. Micrograv vity aasteroid samplee return is not within the sco ope of this stud dy. T The inertial meeasurement is used u for monittoring astronau uts' w weight in micrrogravity envirronment of Intternational Spaace S Station [5]. How wever, inertia measurement must m apply thru ust too the mass beeing measured and thus the system becom mes rrelatively comp plicated. The optical o method ds such as pho oto im maging can observe o the shape s and collor of planetaary ssamples during g sample acquissition. Howeveer, a photo imaage ddoesn't providee sample mass and the viewiing aperture may m bbe obstructed by b dust coverin ng the lens. Maass measuremeent bby spring-masss system appeaars to be the siimplest approaach ffor sample weig ght measuremeent on the planets/moons.

Figg. 1. Cross-secctional view off planetary sample mass sensor moun nted at bottom m of sample caanister.

3. S SENSOR DE ESIGN Figure 1 shows a crooss-sectional vview of the saample mass sensor as it is mounnted at the bottom part off a sample canisterr [6]. It is assuumed the voluume of sample canister is approxiimately 1 literr. Only the boottom section of sample canisterr is shown inn Figure 1. Thhe sensor is pplaced as a false-boottom and so that the weighht of planetaryy sample is measurred during the sample acquiisition processs. The top plate iss a thin stainless steel SUS304 membranne that will deform m by the sampple load. The middle layer is a rigid substraate that has ann array of electrodes on thee top side. There iis a narrow gapp between the S SUS304 top pllate and the substraate allowing thhe deformationn of the top pplate to be capacittively detected.. ICs and otherr discrete compponents for the cappacitance readoout are mounteed on the backkside of the substraate. There is an additional bback plate justt under the substraate. The shape of the two pllates (top and bottom) is identicaal except that the bottom pllate has a cennter hole to allow ttall circuit com mponents protruusion. The SU US304 back plate i s used as a reference cappacitance for calibration purposees. The stack of top plate, suubstrate and boottom plate is attacched to the botttom of canisteer by an Alum minum 6061 retainerr ring and preessure-loaded ffasteners. Smaall gaps on the moounting hole abbsorb coefficieent of thermal expansion (CTE) mismatch whhen the sensoor is exposedd to large temperaature change. A An o-ring placced on the side of retainer ring prrevent small pplanetary sampple particles frrom falling into narrrow gaps on tthe edges of pllates and substtrate as this may caause sample contamination. Air vent holees equalize the cavvity pressure too the ambient pressure. A m microporous

D Deflection of th he spring in su uch spring-masss system can be ddetected by tran nsducers such as a piezoresistiv ve and capacitiive ssensors. Howeever, temperatu ure sensitivity y is an inhereent ddisadvantage in piezoresistiv ve sensors. Fo or the planetaary aapplications th hat undergo large l temperatture change (or ( ccycles), capacitive method appears to have significaant aadvantage. T The SVS senssor described here is an in-situ sensor th hat m measures the saample weight during d sample acquisition. Itt is ddesirable for the t sensor to be integrated d to the samp ple ccanister or cach he. In other wo ords, separate weighing statiion too measure caanister weightt is not an attractive a optiion bbecause it requ uires complicatted maneuver by b a robotic arrm. T The integrated d SVS sensorr will need to be design ned liightweight, sm mall volume and shock tolerant. This is bbecause the sam mple canister is a primary paayload of a retu urn vvehicle that would w endure re-launch fro om planet/moo on, 2

where q is the pressurre, r is the radiius, and h is thee thickness of the ccircular diaphraagm. D is flexuural rigidity;

ffilter is placed on the vent ho ole so that dustt particle will not n innterfere with capacitive c meassurement.

Eh 3 D , 12(1   2 )

T The deformatio on of the SUS S304 top platee is illustrated in F Figure 2. Assu uming a light and a uniform lo oading (Figure 2B B), the deform mation is relaatively large at a the center of m membrane com mpared to the outside. o Thereffore, capacitan nce aat the center (C C_ctr) shows laarger increase compared to the t ccapacitance at the outside (C C_out). When the applied lo oad inncreases, the center c portion of o the membraane will touch the t ffloor (Figure 2-C). 2 The C_cttr is now shortted and becom mes uunable to meassure, but the outside o C_out is still availab ble. U Using a series of concentric distributed capacitors expan nds thhe overall pressure measurem ment range. Figure 2-D show ws thhe membrane is experiencin ng extremely high h load. In th his ccase, the memb brane is mostlly touching an nd stopped at the t uunderlying sub bstrate and th he capacitive readout is not n ffunctional. How wever, the mem mbrane does not n rupture eassily bbecause its meechanical stren ngth is significcantly reinforcced bby the underly ying rigid sub bstrate that preevents excessiive sstrain. Therefore, the sensor can be made highly sensitiive bbut robust agaainst possible shocks that may m be expectted dduring launch and a landing phases.

(2)

where E is the Youngg's modulus, v is the Poisson'ss ratio. Assumiing 0.1kg loadd is uniformlyy applied overr a circular SUS3004 plate with 90mm diam meter (2x5mm= =10mm is  subtraccted from mem mbrane diameeter 100mm ddue to the retainerr ring on the edge), 200m m in thickness, maximum deflectiion is estimateed as 66.8m from equationn (1). Here, the 0.11kg load is eequivalent to 0.27kg samplle load in Martiann gravity and 00.6kg sample looad in lunar grravity. Figure 3 shows thhe deflection of the samee SUS304 membrrane under 0..1kg uniform load as calcculated by ANSYS S finite elemeent analysis simulation. The obtained center deflection 68..6m is reasonnably close too the value calculaated from Timooshenko's equaation (1).

Fiig. 3. ANSYS ssimulation. 0.1kg weight ap pplied unifformly over th he sensor surfface. (z-axis sccaling is exaggerated d) Figure 4 shows electtrode pattern tthat was used in the first sensor prototype. Seeven concentriic electrodes aare placed. Surfacee areas of sevven electrodes are approximaately equal so that the initial capacitance of thee parallel platee capacitors will bee identical. Cappacitance of pparallel plate ccapacitor is given bby,

Fig. 2. Wide range sensing g by distributeed capacitors. (Not to o scale)

C 

T The geometry of the pressu ure-sensing meembrane is veery ssimilar to MEMS M touch mode pressu ure sensors [7]. O Originally baseed on Timosheenko's textboo ok [8], maximu um ddeflection w0 of a circularr membrane at its center is eexpressed as:

qr 4 64D

1 w2 1  0.48 20 h

,



(1) 3



(3)

where  is the permiittivity, A is thhe area and d is the gap distancce. Shield wirres are placedd between eleectrodes in order too reduce RF innterference.

F FEM Analysis

w0 

A , d

substraate. 200 m too 500 m thickk SUS304 topp plates are placed on top of thee glass substraates for testingg. Figure 7 shows the sensor proototype that iss being tested by placing calibrattion weights onn the SUS304 ttop plate. Figure 8 shows a ploot of capacitannce change onn this SVS sensor prototype. Actual thickkness of thee SUS304 membrrane is 206 μm m and the Silicoon spacer thickkness is 94 μm. Neearly uniform load is applieed by placing numerous small weights evennly on the SU US304 membbrane. The overall sensor charaacteristics appeear to be sim milar to the finite eelement simullation result sshown in Figuure 5. The centerm most channel C Ch4 increases the quickest annd the rate of capaacitance changge is approximaately 1pF/gram m when the load is less than 20 grram. The outerr electrodes Chh1 and Ch2 ramps up well abovee the 100 gram m load and ddemonstrate wide-raange sensing caapability of thee mass sensor. However, the ratee of capacitancce increase is sm maller at high--load range (>100ggram) because the center of m membrane is toouching the glass suubstrates. Succh contact modde is not incluuded in the presentt finite elementt model. Otherr deviation from m the finite elemennt model mayy be attributeed to small fabrication deviatioon (membranee thickness 2066 μm as opposed to 200 μm in m model, and inittial gap 94 μm m rather than moodeled 100 μm, meembrane warpaage) and stray capacitance duue to small wire/eleectrode overlapps.

Fig. 4. Electrode pattern of o first proof-o of-principle proto otype. B Based on the finite f element analysis a results, capacitance of thhese channels is estimated an nd plotted in Figure F 5. The gap g w width is assum med as 100 m and the load is appliied uuniformly. Duee to the symmeetry, channels 5, 5 6 and 7 resu ults aare omitted beccause they are identical i to chaannels 3, 2 and d 1, rrespectively. Th he center chan nnel Ch4 ramps up the quickest aand the capacitor become sh horted at abou ut 100 gram lo oad w when the top plate p and electtrode are in co ontact. The ou uter eelectrodes such h as Ch1 ramp ps up relatively y slow and it can c m measure 800 gram loading g or above. Therefore, T inn ner eelectrodes havee higher sensittivity for a smaall loading ran nge w while outer eleectrodes have relatively r low sensitivity s over a w wide range.

Fig. 6. Distributed d electrode patttern on glass substrate.

Fig. 5. Ch hange of capacitance. (FEM M simulation)

4. PR ROOF-OF-P PRINCIPLE E TEST T The first SVS mass sensor prototype is fab bricated by usiing M MEMS fabricaation techniquee. Figure 6 sho ows a picture of thhe prototype substrate. s A 10 00 mm diameteer glass substraate is anodically bonded with th hermally oxidizzed silicon wafer 1100 m in thicckness. The silicon wafer is etched by Deeep R Reactive Ion Ettch (DRIE) to form spacer riing at the edge of ssubstrate. Metaal electrodes arre lift-off patterrned on the glaass

F Fig. 7. Proof-oof-principle sen nsor tested wiith callibration weigghts. 4

power and conform ming to the exxtremely limitted spatial constraaints afforded to it. The sensor electrronics are compriised of the capacitance readout circcuitry, bidirectioonal communiccation to and ffrom the host spacecraft, and poower regulatioon for the aactive circuitrry on the electronnics board. T The following paragraphs wiill describe electronnics integratioon to the senssor itself, the method of deducinng sensor capaacitance and hoow this data is sent to the host sp acecraft for proocessing. The lim mited overall aarea and heightt in which the electronics reside ccreated a spatiaal limitation inn which the dessign would be connstrained. Thhe electronics are mounteed directly beneathh the sensor substrate in oorder to preseerve signal integritty and providee an all-in-onne solution forr the mass sensor itself. The height limitaation placed upon the electronnics (less than an 0.25 inch and falling toowards the edges of the sensoor) greatly hhandicapped ccomponent selectioon and called ffor a design w with minimal ppart counts. The ressulting circuit design was a simple circuiit topology for sennsing the capaccitance of eacch of the sevenn channels and reeporting the values of eaach sensor too the host spacecrraft.

Fig. 8. Ch hange of capaccitance experiimentally measured d on proof-of-principle masss sensor.

N Non-uniform m loading D During the maass measuremeent of solid ro ock and soil, the t ssample may be distributed unevenly u overr the membran ne. W When a 0.1 kg mass is placed d over a 20 mm m diameter circcle aarea at the cen nter, maximum m deformation is 204.6 m as eestimated by ANSYS sim mulation. This deflection is aapproximately three times larger compaared to uniforrm looading of the same s mass. Fig gure 9 shows another a simulatted eexample wheree the same 0.1 kg load is app plied over 20m mm ddiameter circle area that is 20 0 mm off-centeer. The maximu um ddeformation is estimated as 150.9 m an nd the deform med zzone moved offf-center to thee right. Algorrithm is curren ntly bbeing developeed to accurateely estimate saample weight by sstatistical app proach couplled with ro obotic actuatiion m mechanism to tilt the samp ple canister fo or moving solid ssample within the t canister.

The theeory of operation for the senssor electronics is that of a simple current sourcee charging a caapacitor. The seensor itself act likees seven discrette capacitors, w where the top ddeformable plate ffunctions as a common electrode connnected to electriccal ground foor all seven ccapacitors. Thhe positive electrodde for each caapacitor is thenn individually interfaced to the sense circuittry. A microccontroller com mmands an analog multiplexor ((AMUX) to seelect between the seven sensor electrodes, aas well as a temperature ssensor and referennce capacitor for calibratiion purposes. When a specificc sensor electtrode is targetted, a counterr is started within the microcont ntroller and a precision currrent source chargess the capacitorr. The voltage is then monitoored on its output by a fast coomparator. Onnce the capaciitor charge reachess the level oof a precisionn reference vooltage, the comparrator changes sstates and signals the microcoontroller to stop itss running counnt. This count is then used tto decipher the cappacitance of the targeted sensor. This process is repeateed for all sevenn capacitors annd an accuratee model of weight distribution annd overall masss can be obtaiined by the spacecrraft electronicss. The miicrocontroller uutilizes a 16-bit counter thatt ultimately providees the resoluttion and accurracy of the ccapacitance measurrement. When interfaced withh the prototypee sensor as seen inn Figure 7, tthe readout electronics werre able to providee a capacitancee measuremennt with 0.1 pF resolution. The couunter also alloowed for a fulll-range capacittance sense of up too 4 nF with thhe aforementionned resolutionn across the entire rrange. The flooor of the capacitance measurement is limitedd by the strayy capacitancee inherent to the PCB, componnent input cappacitances, andd the wires thhat connect the sennsor to the elecctronics board.. Figure 10 shhows a toplevel bllock diagram oof the electroniics board.

Fig. 9. AN NSYS simulatio on of uneven loading. l 0.1kg g mass (20mm m diameter ciircle footprintt) placed 20mm m off-ceenter. (z-axis scaling s is exag ggerated)

5. CAPAC CITANCE READOUT R T CIRCUIT T The goal of th he SVS capaciitance readout circuitry was to pprovide in-situ u mass sensin ng while conssuming minim mal 5

T The capacitancce value is computed for eaach of the sev ven cchannels of th he sensor witthin the microcontroller. The T m microcontrollerr then outputts this data, in addition to aancillary data such as tempeerature and staatus, to the ho ost sspacecraft usin ng a LVDS trransmitter. A single +5V rail r pprovided by th he spacecraft powers p the bo oard and a lineear vvoltage regulator creates the +3.3V necessaary to operate the t bboard.

R REFERENC CES [1] D. S. Burnett, B. L. Barracclough, R. Beennett, M. Neuugebauer, L. P P. Oldham, C. N N. Sasaki, D. Sevilla, N. Smiith, E. Stansbeery, D. Sweettnam, R. C. W Wiens "The Gennesis Discoverry Mission: Reeturn of Solarr Matter to Eart rth", Space Sciience Reviews, Vol. 105, ppp. 509-534, 20003. [2] D. E E. Brownlee, P P. Tsou, J. D. Anderson, M. S. Hanner, R. L L. Newburn, Z Z. Sekanina, B. C. Clark, F. H Horz, M. E. Zolenski, J. Kisseel, J. A. M. MccDonnell, S. A. Sandford, A. J. Tuzzolino, "Stardust: Com met and intersstellar dust sam mple return misssion", J. of G Geophys. Res.,, Vol. 108, E100, 8111, doi:10..1029/2003JE0002087, 2003. [3] J. Kawaguchi, H H. Kuninaka,, Fujiwara, T T. Uesugi, "MU USES-C, Its laaunch and earlly orbit operatiions", Acta doi: Astr tronautica 559, pp. 6669-678, 2006, 10.11016/j.actaastroo.2005.07.002 [4] D. A A. Papanastassiou, G. J. Waasserburg, D. S. Burnett, "Rbb-Sr ages of luunar rocks from m the sea of trranquillity", Eart rth and Planetarry Science Lettters, Vol. 8, Issuue 1, pp. 119, 1970 doi:10.10016/0012-821X X(70)90093-2

p-level block diagram d of SV VS electronics.. Fig. 10. Top

[5] J. J. Uri, C. P. Haven, "Accomplishhments in bioaastronautics rresearch aboaard Internationnal Space Stattion", Acta A Astronautica, 556, pp. 883-8889, 2005. doi::10.1016/j.actaaastro.2005.01.0014

6. SUM MMARY A proof-of-prrinciple samp ple verificatiion system is ddeveloped. Thee objective of the t sensor is to o perform in-situ m measurement of o planetary rock/soil r samp ple mass duriing ssample accumu ulation. The sensor s is desiigned to preveent ssample contam mination, to withstand larrge temperatu ure cchange, to miinimize weigh ht and volum me, and to haave rrobustness agaiinst shock. The proof-of-prin nciple sensor has h sshown high sen nsitivity of 1pF F/gram at low mass m range wh hile m maintaining caapability to measure m largee mass. Furth her ddevelopment will w include buiilding algorith hm for estimatiing uuneven sample loading, and use u of PCB-baased substrate for f ccircuit integrattion. Also, en nvironmental testing t (therm malvvac, shock, etcc.) will be performed for further fu validatiion tooward future planetary p sample return missions.

[6] JPL L New Technollogy Report #477690. [7] W. H. Ko, Q. W Wang, "Touch m mode capacitivve pressure senssors", Sensors aand Actuators 775 (1999), pp. 242-251. [8] S. P. Timoshenkko, S. Woinow wski-Kreiger ""Theory of Plattes and Shells",, McGraw-Hilll, 1940.

ACKNOWL A LEDGEMEN NT T The authors would like to th hank Mr. Don Hunter and Mr. M E Eduardo Urgilees for valuable suggestions. This T research was w ccarried out at the Jet Prop pulsion Laboraatory, Californ nia IInstitute of Tecchnology, undeer a contract with w the Nation nal A Aeronautics and Space Admin nistration.

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mart power an on-ggoing collaborrative effort onn developing sm integraated mixed-signnal electronicss for sensors, actuators, and poower manageement and diistribution sysstems. Dr. Mojarrradi has over 220 years of tecchnical, managgement and academ mic experiencee. He receivedd his Ph.D. inn electrical engineeering from thee University off California, Los Angeles (UCLA) A) in 1986. He has tweenty-two pateents, forty publicaations and is a member of IE EEE. Prior to jjoining Jet Propuls lsion Laboratorry, Pasadena, CA, he was ann Associate Professsor at Washinggton State Uniiversity and thee Manager of the mixed-voltagee/specialty inteegrated circuit group at the Xerrox Microelectrronics Center, El Segundo, C CA.

BIOGR RAPHY Risaku Toda (Member, IE EEE) received d a B.S. degree in n Applied Physsics and a Ph..D. degree in mechatronics m and precisiion engineering from fr Tohoku University, U Send dai Japan. He allso received a M.S. degree in materials sciience and en ngineering fro om University of o California Los Angeles ((UCLA). Priorr to joining Jet J Propulsion Laboratory, he w worked for Fo ord Motor Co ompany Japan n Ltd. and Ball B SSemiconductor Inc. in Allen TX. Since 20 003, he has beeen w with Nano and Micro Syystems group at NASA Jet J P Propulsion La aboratory. Hiis research interests inclu ude sensors, fa fabrication of MEMS actuators and a nnanotechnologyy devices forr aerospace applications. a He H hholds several US U patents.

Ash hitey Trebi-Olllennu (Seniorr Member, IEE EE) is a techhnical group lleader and robbotics softwarre engineer IIV of the Mobbility and Mannipulation grouup at NASA Jet Propulsion Laboratory, California Insttitute of Technoology, where hhe has been sincce 1999. He reeceived his Phh.D. degree in control system ms from Royaal Military Cranfield Univversity, Unitedd Kingdom Collegee of Science, C in 19996 and B.Eng.. (Hons) from m Queen Mary ry College, Univers rsity of Londonn, United Kinggdom in 1991. He was a researcch scholar aat Institute of Complex E Engineered Systemss, Carnegie M Mellon Universsity from 1997 to 1999. His rresearch intterest includdes planetaryy rovers, manipuulation, distribbuted mobile roobotics, sensorr networks, reconfig igurable roboots, mobile roobot architecttures, and man–m machine interacction and has ppublished wideely in these areas. Dr. Trebi-O Ollennu receiived the 20008 NASA Exceptiional Engineeering Achievvement Medall for his contribbutions to the M Mars Explorattion Rover misssion, 2007 Outstannding Engineer Award from IEEE Region 66, 2007 Sir Monty Finniston Acchievement Meedal from Insstitution of Engineeering and Teechnology, U.K K., and 2010 Specialist Silver A Award from the Royal Aeronnautical Societyy, U.K. Dr. Trebi-O Ollennu is a F Fellow of the Innstitution of E Engineering and Te Technology, U U. K., a Felloow Royal Aeeronautical Societyy, U.K. and a Senior Membber of the IEE EE. He has severedd as a Guestt Editor of tthe IEEE Robbotics and Automaation Society M Magazine.

Colin McKiinney received d his B.S. degrree in Electrica al Engineering g from Cal Po oly, San Luis Ob bispo in 2008 and is curren ntly pursuing hiss M.S.E.E. at the t University of Southern Ca alifornia whilee employed fu ulltime at the Jet J Propulsion n Laboratory. He H is a membeer of the Adva anced Instrumeent E Electronics gro oup under Moh hammad Mojarrradi at JPL with w a primary foccus of analog and mixed-siignal electronics ddesign. He is also active in th he research and d developmentt of thhe enabling technologiees for low--noise extrem me eenvironment (wide-tempeerature, hig gh radiatio on) innstrumentation n-quality circuit design and im mplementation n. Shannon P.. Jackson receeived his B.S. in Electrical Engineering E fro om the Californ nia Polytechnic University, Pomona P and has h been at NAS SA’s Jet Propulsion Laborato ory (JPL) sincce 1985. He H has beeen instrumentall in developiing space flig ght hardware th hat has flown on Earth orrbit aand Mars misssions, as welll as developiing and fieldiing nneutrino detecctors in Anta arctica. His interests i inclu ude m micro-power applications of micropro ocessors, sma art ssensors, fiber optics, o and wireeless systems. He holds paten nts inn the area of signal processin ng and wirelesss Sensor-Webss.

Hariish M. Manoohara (Membeer, IEEE) receiived the Bachelor of E Engineering degrree in instrumeentation technoology from the B Bangalore Unniversity, Indiaa in 1989, folloowed by an M M.S. degree iin nuclear enginneering in 19992, and a Ph.D D. degree in enginneering sciennce in 1997 from the Louisiaana State Univeersity, Baton R Rouge, Louisianna. He joinned the Centeer for Advancced Microstrucctures and Devicess (CAMD) at the Louisiana State Univeersity as a Researc rch Associate inn January of 19997. In 1998, he became the Asssistant Professsor of Reseaarch at CAMD D, and an Adjuncct Faculty Mem mber of the D Department off Electrical Engineeering. Whilee at CAMD, hhe developed new X-ray microfa fabrication tecchniques. In November off 2000 he joined the Jet Propuulsion Laboratoory to developp advanced

Mohammad d Mojarradi (Senior (S Member, IEEE) is a Principal att Jet Propulsiion Laboratory in Pasadena a, CA. A Seniior Member off IEEE, he wass the chair of the t F Valleey Section and an IEEE San Fernando integrated circuit dessign specialiist. Specificallyy he is an expeert in developiing mixed-signa al/mixed-voltag ge electron nic ccircuits for sm mart power systems, s sensors and micrrom machined elecctromechanical interface applications. a He H leeads a consorttium of universsities developin ng electronics for f eextreme environments and manages m the deevelopment of the t eelectronic cirrcuits for th he "thermal cycle resista ant eelectronics" tassk for Mars Sccience Laborattory. He also has h 7

silicon and vacuum electronic devices for THz applications. Since 2005, he has been leading the Nano and Micro Systems (NAMS) group at JPL and has developed carbon nanotube field emitters, nanoelectronic devices, miniature spectroscopic instruments, and MEMS for space, defense, medical, and commercial applications.

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