SMART1 operations experience and lessons learnt

Acta Astronautica 61 (2007) 203–222www.elsevier.com/locate/actaastro

SMART-1 operations experience and lessons learnt

Octavio Caminoa,∗, Maria Alonsob, Daniel Gestalc, Jurriaan de Bruind, Peter Rathsmane,Joakim Kugelberge, Per Bodine, Sascha Rickenc, Rick Blakeb, Pablo Pardo Vossd,

Luca Stagnarof

aESA/ESOC, Robert-Bosch Stasse 5, Darmstadt, GermanybScience Systems GmbH, Darmstadt, Germany

cLSE Space Engineering and Operations AG, Darmstadt, GermanydVega GmbH, Darmstadt, Germany

eSwedish Space Corporation, Solna, Swedenf ESA/ESTEC, Noordwijk, The Netherlands

Available online 27 March 2007

Abstract

SMART-1 is the first of a series of ESA Small Missions for Advance Research and Technology where elements of the platformand the payload technology have been conceived as a demonstration for future cornerstone missions and an early opportunity forscience. SMART-1 has also been an opportunity to experiment with new ways of conducting ground operations taking advantageof both increased satellite autonomy and ground automation tools. The paper will focus on three areas:

1. The accumulated performance of the technology demonstration components since launch as the electrical propulsion engine,the triple-junction solar cells, the lithium-ion batteries, the 32 bit CPU ERC32 Single Chip, the CAN bus, the DTU StarTrackers and the complex on-board autonomy.

2. The changes implemented on-board and on the ground during the lunar phase to increase the data return.3. The pros and contras in some of the choices made for SMART-1, the developments and solutions implemented to mitigate

the problems, the tools developed to automate the operations and the distribution of data.

© 2007 Elsevier Ltd. All rights reserved.

1. SMART-1 mission

The satellite, manufactured by the Swedish SpaceCorporation (SSC) as prime contractor, was launched

∗ Corresponding author.E-mail addresses: [email protected] (O. Camino),

[email protected] (M. Alonso), [email protected](D. Gestal), [email protected] (J. de Bruin),[email protected] (P. Rathsman), [email protected] (J.Kugelberg), [email protected] (P. Bodin), [email protected] (S.Ricken), [email protected] (R. Blake), [email protected](P.P. Voss), [email protected] (L. Stagnaro).

0094-5765/$ - see front matter © 2007 Elsevier Ltd. All rights reserved.doi:10.1016/j.actaastro.2007.01.042

on 27th of September 2003 and spiralled out over a14-month period until being captured by the Moonon November 15th 2004, thus successfully achievingthe primary objective set to demonstrate solar electricpropulsion.

The initial orbit target was optimized in two phasestaking advantage of the good performance of the elec-tric propulsion (EP). The first brought the orbit apolunedown to 3000 km from the initial 10 000 km; the seconddone during August and September 2005 optimized theorbit visibility for science increasing the argument ofpericentre.

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204 O. Camino et al. / Acta Astronautica 61 (2007) 203–222

The science experiments on-board SMART-1 were:

• EPDP to monitor the working of the propulsion sys-tem and its effects on the spacecraft.

• SPEDE to also monitor the effect of the propulsionsystem and to investigate the electrical environmentof the Earth–Moon space.

• KaTE to test new generation of communication tech-niques with Earth.

• RSIS uses the KaTE and AMIE instruments to inves-tigate the way the Moon wobbles.

• OBAN is a software to allow the spacecraft to guideitself to the Moon.

• AMIE to test a miniaturized camera and take colorimages of the Moon surface.

• SIR to search for ice and make a mineralogical map-ping of the Moon.

• D-CIXS to investigate the composition of the Moon.• XSM to calibrate the D-CIXS data and study solar

X-ray emission

1.1. Mission phases

◦ Launch and early orbit phase. Launch on September27th 2003, initial orbit 7029 × 42 263 km.

◦ Van Allen belt escape. Continuous thrust strategyto quickly raise the perigee radius. Escape phasecompleted by December 22nd 2003, orbit 20 000 ×63 427 km.

◦ Earth escape cruise. Thrust around perigee only toraise the apogee radius.

◦ Moon resonances and capture. Trajectory assistsby means of Moon resonances. Moon capture onNovember 15th 2004 at 310,000 km from the Earthand 90,000 km from the Moon.

◦ Lunar descent. Thrust used to lower the orbit, opera-tional orbit 2200 × 4600 km.

◦ Lunar science. Until the end of lifetime in September2006, interrupted only by a one-month reboost phasein September 2005 to optimize the lunar orbit.

◦ Orbit re-boost phase in June/July 2006 using theattitude thrusters to adjust the impact date andtime.

2. Launch and early orbit phase (LEOP) and VanAllen belt escape

LEOP starts at launch vehicle separation, and wasdedicated to the commissioning of all the critical plat-form units before EP orbit raising could commence.This was achieved with two teams working 12-h shiftsaround the clock. Platform commissioning was ex-

ecuted rapidly to allow the EP orbit raising phaseto start as soon as possible, thus limiting radiationexposure.

The major concern after launch during the first threemonths of the mission was to leave the radiation beltsas soon as possible in order to minimize the degra-dation of the solar arrays and the Star Tracker CCD.The team was confronted with operational challengesof varying degrees, further aggravated by the strongestsolar flares occurred in 40 years all while the spacecraftwas in continuous thrust spiralling out of the radiationsbelts.

The first and most critical problem came after thefirst revolution when a failure in the on-board error de-tection and correction (EDAC) algorithm forced an au-tonomous switch to the redundant on-board computer.The analysis of the spacecraft telemetry pointed directlyto a radiation-triggered problem with the EDAC inter-rupt routine [1].

Other anomalies during this period were a combina-tion of environmental problems: high radiation doses,especially in the start trackers and on-board softwareanomalies. The Reed Solomon encoding became cor-rupt after switching data rates; this on-board problemdetected during the LEOP phase had to be overcome byprocedures and changes on ground.

The Star Trackers were also subject of software tun-ing and adjustments during the earth escape and werecause of some of the EP interruptions [2].

The EP showed sensitivity to radiation inducing shut-downs. This phenomenon identified as the optocouplersingle event transient (OSET), initially seen in LEOPduring the first firing using cathode B, was characterizedby a rapid drop in anode current triggering the alarm‘flame out’ bit causing the shutdown of the EP.

A joint analysis done by ESOC/ESTEC, Snecmaand ETCA experts showed the problem to be radiationinduced optocoupler sensitivity. The recovery of suchevents was to restart the thruster. This was manuallydone until an on-board software patch (OBSW) wasdeveloped to automatically detect it and initiate anautonomous thruster restart. The consequences werereduced to a minimum with the only operational im-pact being in the orbit prediction calculation used forthe ground stations to track the spacecraft.

The different kind of anomalies caused frequent inter-ruptions in the desired continuous thrust of the EP (seeFig. 1). This led to an increase of the ground stationssupport and overtime of the flight operations team whohad to quickly react to the interruptions. Their recoverywas sometimes time consuming, especially when thespacecraft was found in SAFE mode.

O. Camino et al. / Acta Astronautica 61 (2007) 203–222 205

22%(10)

4%(2)

4%(2)

15%(7)

55%(38)

AOC Fine Pointing lossOSET (Flame Out) (Re-boost phase not included)Thermothrottle current fluctuationsPower Processing Unit (PPU) Indicated OFFSpacecraft in Safe Mode

Fig. 1. Spacecraft events that caused unexpected EP power off.

3. Earth escape and transfer to the Moon

The Van Allen belts escape phase used a continuousthrust strategy, initially thrusting along the velocity vec-tor. This strategy was chosen to quickly raise the perigeeradius above a target 20,000 km. At this point, whichwas reached on January 7th 2004, the apogee radiuswas increased to 57,500 km.

Close to the end of the Van Allen belt escape phasethe thrust strategy was changed to thrust perpendicularto the position vector (in plane). In mid March 2004, theeclipse season was such that the spacecraft was passingthrough the Earth’s shadow near the apogee. Eclipsesat this altitude put at prove the battery capacity for 2 hand 15 min. Eclipse times in excess of this value wereactively avoided by performing a rotation of the line ofapsides, so that they occurred sufficiently far away fromapogee to meet the constraint. Later on, the orbit wasrotated back to the ecliptic to properly reach the Moon.

The strategy to achieve these changes was based ona Hohmann transfer. Up to about mid September thrustarcs near perigee pump up the apogee to a radius near300,000 km. Subsequently, apogee thrust arcs increasethe perigee to a radius of about 170,000 km. The apogeethrust arcs were tilted out of plane to rotate the orbitalplane to the final inclination of about 12◦ (see Fig. 2).

The apogee had to be increased from 67,500to 300,000 km, the perigee from 20,700 to about160,000 km and the inclination w.r.t. the orbit planeof the Moon reduced from 33◦ to 12◦. Thanks to thefavourable launch date, the argument of perigee w.r.t.the orbit plane of the Moon had already the requiredvalue of 178◦.

Fig. 2. Pre-launch planned strategy until Moon capture.

The transfer phase was used to do a more exhaustivecommissioning of the instruments and to commencescientific observations although this was not part of theoriginal mission operations plan.

3.1. The resonances

A considerable part of the perigee raising was ob-tained by including Moon resonances. From an apogeeradius of 200,000 km known onwards, the Moon startsto significantly perturb the orbit when it is in the vicin-ity of the spacecraft at apogee. If the Moon is ahead ofthe S/C, the Moon will exercise a perturbing force onthe S/C in velocity direction. This results in an increaseof the perigee height. If the Moon is behind the S/C,the perigee is reduced. An optimal perigee increase isobtained when the Moon is about 15◦ behind the S/Cwhen the S/C is at apogee. This angle can be controlledby adjusting the orbit periods prior to the resonance.

Initially, it was foreseen to perform three such res-onances and three lunar swings-by, before the capturemanoeuvre. Analysis of the on-board propellant andpower budgets around the time of the planned lunar cap-ture allowed a new strategy to be developed. The threelunar swings-by were removed, which had the benefitsof reducing transfer time, and the time spent around theoperationally critical weak stability boundary betweenthe Earth and the Moon.

The first and weakest resonance was at an apogeeradius of 230,000 km on August 19th 2004. The secondwas stronger and it occurred on September 15th 2004,four revolutions later of the first, at an apogee radiusof 290,000 km. Finally the strongest resonance was onOctober 12th 2004, three revolutions later of the second


Fig. 3. Moon transfer orbit. Resonances and Moon capture.

at an apogee radius of 324,000 km. The three Moonresonances were about 27 days apart and their effectwas to raise the perigee and to rotate the orbit both ininclination and argument of perigee [3].

On the 15th November 2004, SMART-1 became thefirst EP mission to escape Earth orbit with the use ofEP, the first to use EP to enter into orbit around anothercelestial body, and Europe’s first lunar mission. At thepoint defined as capture, the spacecraft passed through aposition 310,000 km from the Earth and 90,000 km fromthe Moon in free drift. To achieve this, the EP thrusterhad been started 288 times, accumulating 3652 h firingtime, since launch vehicle injection into GTO (Fig. 3).

The main problems during this phase of the missionwere caused by the numerous and highly constrainedattitude profiles [4].

3.2. The Star Tracker high temperatures

The cameras of the SMART-1 Star Trackers were de-signed to withstand BOL (beginning of life tempera-tures) between −55 and 20 ◦C. A radiation dose of about550 rad on the cameras based on an expected radiationenvironment within the Van Allen radiation belt wouldallow for temperatures only up to 0 ◦C. This is becauseof an exponential expressivity of the radiation inducedhotspots to the CCD.

During October 2003, the FCT realized that the StarTrackers cameras were working most of the time at

temperatures above 10 ◦C. Moreover, since the radiationdoses received were higher than expected, the operatingtemperature of the cameras was even reduced furtherto 7–8 ◦C. Camera 2, for instance, was working 2/3 ofthe orbit at temperatures above 10 ◦C and showing peaktemperatures close to 18 ◦C.

The combination of high temperatures and high ra-diation doses on the CCD of the cameras degraded ingeneral the performance of the CHUs, especially theperformance of CHU2, and thus it reduced the redun-dancy of the Star Trackers. This increased the chancesof loosing fine pointing control which causes the shutdown of the EP engine.

In addition, the combined effect of high radiationdoses and high temperatures impacted the removal ofstray light effects which caused an overload of the dataprocessing unit of the Star Trackers. All the above ledto request DTU to develop a software patch for the StarTrackers in order to increase the operating temperatureof the CHUs and reduce the processor load of the CHUs.Until the software patch was ready, the FCT had tochange the star detection thresholds in order to allowmore stars to be detected.

Apart of the abovementioned Star Tracker attitudeinvalidation due to high temperatures, other problemsaffected their performance like the rich star field, theprocessor overload or the time synchronization [5].

Fig. 4 shows the spectacular improvement of the StarTracker performance (reduction of invalid attitudes for


Fig. 4. Percentage of invalid attitude per Star Tracker camera per day.

Fig. 5. EP power set adjustment to available solar array generation power.

CHU2 and CHU1) after the installation of the new StarTrackers patch.

The EP operations were adapted during this periodto the changing solar flux induced by the seasonal vari-ation of the Earth to Sun distance [6]. Fig. 5 shows themaximum possible EPS thrusting levels that were dueto seasonal solar flux variation.

The EP worked very reliable during this phase. Itsperformance was close to 100% with minor variations(see [7]). The software patch implemented to automat-ically recover from the OSET permitted relaxing the

manning required for the passes. The ground stationsupport was kept at relatively high level surpassing the110 h coverage per week in average. This not only per-mitted to quickly react to anomalies but to dump plat-form data more frequently optimizing the data requiredfor orbit determination.

4. Reaching the Moon operational orbit

The moon spiral commenced immediately after theMoon capture on 15th of November. The apolune height


Fig. 6. Moon spiral orbit (thrust: dotted line; no thrust: continuous line).

of the Moon operational orbit was reduced from theoriginally planned 10,000 km to 3000 km at a cost of5.5 kg of fuel to improve science observations. The orbitperiod was tuned such that each five revolutions thespacecraft was seen by the same ground station.

The R.A.A.N. of the Moon operational orbit (orien-tation of the orbit plane around the Moon pole) wasbiased by −59◦ from the fuel optimal value at a fuelcost of about 1.5 kg). This bias was introduced to shiftthe Moon eclipse period towards the end of the oper-ational period (April 7th–June 21st 2005). The eclipseperiod was finished by the time that EP arcs would beneeded for extending the mission (perilune rising). An-other advantage of this bias was that the Moon capturewas guaranteed in case of EP problems.

The moon operational orbit was achieved by the endof February 2005 requiring another 236 trusts adding953 h to the cumulative time usage of the engine(Fig. 6).

The major concerns faced during phase were:

• The characterization of the Star Trackers camerasthat were expected to be simultaneously above thethermal operational conditions during the descent.

• The design of the on-board stores that was not de-signed to cope with the volume of data since the or-bit duration had changed from 15 h to about 5, thusmultiplying the data collection and the amount ofcommanding well above the initial estimations.

All actions assigned during the Lunar OperationsReadiness review held on November could be com-pleted by the time of the arrival to the Moon opera-tional orbit. Among the changes implemented, therewas a significant on-board software change that in-cluded a complete redesign of the payload stores andits operations from ground.

5. In Moon operational orbit. The scientificmission

The arrival to the Moon operational orbit implieda change in the primary objective of SMART-1.After having achieved the primary goal of testingand validating the EP, the mission was handed overto the Solar Systems Science Operations Divisionthat put a new primary objective: ‘perform lunarscience’.

This phase commenced in March 2005 and it had ma-jor consequences in some of the ground systems. Themission planning system had to be upgraded to copewith the increase of command generation and planning.This ultimately led to the automation of more than 95%of the command generation during the year through con-tinuous upgrades to the system.

The science operations were paralleled with thepreparation of the mission extension taking advantageof the remaining Xenon estimated to be in the order of6 kg.


5.1. SMART-1 mission planning system (MPS)

The principle mode of operation was that all rou-tine platform and payload operations were pre-plannedand executed according to the flight operations plan(FOP), driven by on-board time tagged commands (max10 000) that were up-loaded during each ground stationpass by the FCT.

The purpose of the SMART-1 MPS was to ensure theoverall consistency of these operations request againsta variety of operational and spacecraft resource con-straints, such as power, storage, downlink capacity, priorto generating the required telecommand stack for uplinkvia the mission control system (MCS).

The volume of commanding for two of the instru-ments, the AMIE and EPDP payloads massively ex-ceeded by seven times the level for which the space andground segments were designed.

The problem was temporarily resolved by a reductionin the number of AMIE operations per orbit and by areduction in the number of commands per imaging cyclefacilitated by a new issue of the AMIE telecommanddatabase.

To manage the big volume of data generated, the so-lution proposed for the payload stores was an enhance-ment to the mechanism for management of the on-boardsolid state mass memory used to hold platform and pay-load telemetry. The concept based on a cyclic payloadsuperstore for all payload data with a special payloadstore for high priority/high value payload data. In addi-tion, the superstore concept allowed much finer control

Fig. 7. Operational orbit selected for mission extension. (Distance from the Moon centre).

using the MPS over download activities in that down-load could be started/stopped at any time. This allowedthe best possible use to be made of whatever high ratecoverage was available.

The required payload superstore concept was in-stalled on day 2005.040 and performed very well, elim-inating the problem of payload download scheduling.

6. Mission extension, the reboost

The mission extension had not only the objective ofimpeding the spacecraft impact on the Moon in Septem-ber 2005 but the optimization of the orbit for science.The final version was presented by ESOC Flight Dy-namics on May 18th 2005 (Fig. 7). The strategy con-sisted of using all the usable Xenon (4.2 kg out of 6 kgavailable in the tank). It also foresaw a possible utiliza-tion of the remnant 1.8 kg (theoretically not usable). Thetarget was to increase the argument or perilune since theeffect of the earth will cause its decrease an ultimatelythe impact on the Moon in 2006.

After several months of simulation and developmentof special procedures, the thrust spiral started on Au-gust 2nd 2005. The procedures proved successful andpermitted the use of all the possible Xenon down to∼ 280 g at which point the pressure was not enoughto keep the EP in operation causing almost immediateflame-outs after its ignition. The entire reboost consistedof 207 revolutions of about 5 h each and from October1st 2005 the EP was definitively switched off and thescience activities were resumed.


Fig. 8. Part of the reboost strategy. Series of reaction wheel off-loadings that transfer spacecraft angular momentum between the spacecraft+Y and −Y axes. Generates a small �V in the spacecraft +X or −X axes. In between the spacecraft will be slewed such that the �V isaligned with the flight direction.

7. Avoiding SMART-1 impact on the dark side ofthe Moon

The orbit predictions carried early in 2006 indicatedthat SMART-1 impact would occur in the dark side ofthe Moon by mid August 2006. The scientific commu-nity requested ESOC and SSC to investigate ways togenerate a certain delta V that could bring the impacttowards the visible part. This happened at a stage whenall the Xenon gas of the EP had been exhausted. Thesolution was to use the thrusters of the attitude control(see [8]), which were not foreseen for this purpose, ina series of complex maneuvers during 65 orbit revolu-tions implying 520 reaction wheels off-loading maneu-vers (see Fig. 8). They were done in the second halfof 2006 leading the spacecraft to impact on the desiredtarget in the lake of Excellence on the 3rd of September(see [8]).

8. SMART-1 technology demonstrationperformance and lesson learnt

The main areas covered by the technology demon-stration are listed below:

• The EP.• The triple-junction cells solar.• The lithium ion batteries.• The new 32 bit CPU ERC32 Single Chip used by

ESA for the second time after Proba-1.

Fig. 9. SMART-1 Electrical Propulsion thruster (SnecmaPPS-1350-G).

• The CAN bus, A COTS product derived from theautomobile industry.

• The latest generation of the Danish DTU Star Track-ers.

• The next generation of on-board autonomy concepts.

8.1. Electric propulsion

The EP system (SMART-1 primary propulsion sys-tem) was the PPS-1350-G Hall Effect thruster devel-oped by Snecma (Fig. 9), primarily for North SouthStation Keeping of Geo-stationary satellites. Two


Fig. 10. EP efficiency calculated from FD tracking data.

Table 1EP in numbers

EP in numbers

Number of pulses 844Total number of hours fired 4958.3 hFirst pulse 30/09/03 12:25Last pulse 17/09/05 18:45Cathode A firing time 3865Cathode B firing time 1106BB valves activations cycles 1.256.505Xenon at BOL 82.5 kgRemaining Xenon ∼ 280 gEP power set range used duringthe mission

From 649 to 1417 W

Number of OSET 38Total of flight operation proce-dures created

45

Total of EP sequences created 98

important characteristics in its design were the possi-bility to operate over a range of engine power levelsand the EP gimbal mechanism that provides two-axisre-orientation capability for the EP thruster.

The system consists of four main sections: the Xenonsystem, the electrical power system and thruster, the dig-ital interface and communication system and the gimbalmechanism also called EPMEC.

The Xenon was stored in the main Xenon Tank,82.5 kg at launch, under high pressure (150 bar). Adevice called the Bang-Bang Pressure Regulation Unit(BPRU) regulated the Xenon down to a constant low

pressure around 2 bar. This low-pressure Xenon wasthen fed into the adjustable flow regulator, called theXenon flow controller (XFC) that provided fine controlof Xenon mass flow rate in order to keep the anodecurrent, thrust and power constant (see [9]).

8.1.1. EP performanceThe total performance of the EP system at the end of

its life on September 17th 2005 is depicted in Table 1.The nominal number of valve cycles of 1 million

has been exceeded by more than 25% and they did notpresent any signal of degradation in terms of leakage.The overall efficiency was close to 100% during thewhole mission. It only dropped when the remainingXenon gas was unable to provide the necessary pressuredue to extreme operating conditions (Fig. 10).

8.1.2. EP operations experience and lesson learntSMART-1’s primary mission goal has been success-

fully achieved. The extension of the mission and theachievement of a better observation orbit were madepossible thanks to the low degradation of the solar cellsafter crossing the Van Allen belts. This allowed operat-ing the EP sub-system powered by a Hall Effect thrusterat a higher power level.

A set of advanced low pressure EP operations enabledthe residual Xenon mass to be reduced from 1.8 kg toonly ∼ 280 g.


8.1.3. EP lessons learntEP hardware surveillance

◦ SMART-1 has a power hardware surveillance thatshuts down EP after 120s of battery discharge. Ex-ample: heater power peaks can be absorbed by bat-teries; in consequence, EP power utilization could beincreased.

EPDP and SPEDE

◦ SMART-1 has two plasma instruments on-board—availability of their data can be very useful for trou-bleshooting anomalies and thrust level changes.

◦ It is advisable for future missions to carry comple-mentary plasma and EP devices equivalent to theSMART-1 EPDP and SPEDE.

EP OBSW

◦ All surveillances that result in an action (e.g. in EPcore) should include filtering and the filtering shouldbe tuneable by telecommand.

◦ Software surveillances are sufficient monitoring theEP performance and the Plasma.

Patchable EPS Software

◦ EP systems have software loops running to control thethruster. On SMART-1 they are not patchable—sometuneable parameters do provide flexibility. Just to in-dicate its impact, if the SMART-1 PPU software werepatchable the OSET problem could have been solvedat subsystem level. This solution would have beenmore efficient sending ignition pulse immediately toignition electrode instead of lasting 30 min.

A lesson learnt, generic to all systems, deals with the im-munity to the single events transients for which, groundtests are not really feasible at system level, and thusshould imply a specific deep analysis on the S.E.T. con-sequences in the system.

Other suggestions for future EP systems have beenderived from the SMART-1 experience: dynamic powerset to maximize the use available power to increasethrust and/or specific impulse, new algorithms for theutilization of the batteries during short eclipses, alter-native simple ground station systems based on carriertracking and use of data relay satellites (see [10]).

8.2. Attitude orbit control subsystem (AOCS)

The SMART-1 attitude and orbit control system(AOCS) were developed by the SSC (see [11]). The

AOCS uses Star Trackers and reaction wheels as itsprimary sensors and actuators. The AOCS softwaremodule is programmed in MATLAB/Simulink fromwhich C-code is automatically generated. The gener-ated code was then compiled and linked into the generalon-board software. A major AOCS design driver hasbeen the need to accommodate the Electric PropulsionSystem (EPS).

The SMART-1 AOCS design solves the problem ofa changing centre of mass as the EP thruster consumesfuel by using the gimballed mounting mechanism (EP-MEC) of the EP thruster in closed-loop to continuouslyalign the thrust vector with the centre of mass. Thisdesign saves hydrazine mass compared to the useof hydrazine thrusters to compensate for EP thrustmisalignment.

The attitude and rate estimation module estimates theon-board attitude, body rates and sun sensor based sunvector. The module also calculates the reference attitudeand inertial sun direction from the up-linked Chebychevcoefficients. The AOCS FDIR functions consist of unitlevel detection as well as several functions for detectionof higher level anomalies.

The SMART-1 AOCS has a high degree of autonomythat minimizes ground intervention. The initializationof any AOCS mode includes the automatic configura-tion of the necessary hardware and the automatic pro-cesses or functions necessary to achieve the objective ofthe mode.

The AOCS has four modes and several sub-modes:Detumble mode may be entered during any mission

phase and is the entrance mode after launcher sep-aration. It reduces high body rates to an acceptablelevel.

Safe mode is the base mode for all other modes andthe recovery mode for emergency situations. Its purposeis to ensure power generation and ground communica-tion using a minimum of on-board resources. The pur-pose of safe mode fine pointing control is to keep thespacecraft pointing with the EP thruster towards the sunand slowly rotating about this vector.

Science mode is a scientific observation mode usedfor surface and cruise science.

EP control mode is the mode used for firing the EPthruster to modify the orbit.

8.2.1. AOCS technology—sensors and actuators◦ Two Star Trackers (ST). The ST hardware consisted

of a data processing unit (DPU) and two camera headunits (CHUs) connected to it by cable. SMART-1had a nominal and a backup DPU and two CHUscross-strapped to both data processing units. The


Fig. 11. S/C and its coordinates system.

active DPU operated both camera-head units simul-taneously within the avionics design of SMART-1.

◦ Five solid state quartz single-axis angular rate sen-sors. They serve as rate sensors, though in higherspacecraft modes signal integration was performedfor temporary periods of ST loss. The saturation limitfor the angular rate detection was 10◦/s. The accu-racy of these ARS were of the order of 10–30 mdeg/s,but a long-term drift of the order 100 mdeg/s was ex-pected during the mission.

◦ Three coarse sun sensors. They were used for space-craft and solar array pointing in safe mode, and forSun/eclipse detection. The Sun sensors were locatedon the spacecraft. Each sensor is able to cover nearly2� steradians.

◦ Four reaction wheels. They were oriented in a pyra-mid configuration, only three were used simultane-ously to minimize power consumption and providecold redundancy. Maximum angular momentum ca-pability of one wheel was 4 N m s.

◦ The solar array drive mechanism (short name:BAPTA). They allowed the solar arrays to rotatearound the spacecraft Y-axis (see Fig. 11). In theorbit-raising phase this allows the thrust vector andsolar arrays to be optimally pointed at the same time.

◦ The hydrazine monopropellant propulsion subsystem.It consisted of one diaphragm-type propellant tank,two service valves, one system filter, one latch valve,one pressure transducer, and eight 1.0 N monopro-pellant hydrazine thrusters. The thrusters were con-figured and controlled as two redundant branches,each with four thrusters and connected to their cor-responding remote terminal units. The eight thrusterswere mounted onto four brackets that were mountedon the spacecraft −Z side.

◦ The electric propulsion gimbal mechanism (shortname: EPMEC). It was a software-controlled mecha-

Table 2AOCS anomalies caused by hardware (H/W), software (S/W), ther-mal (TH) or radiation environment (RD) problems

Sensor Anomalies related to. . .

H/W S/W TH RD

Star Tracker – 3 1 1Sun sensors – – – –Gyros – 1 – –Reaction wheels – – – –Hydrazine S/S – – 1 –S/A drive mechanisms – – – –EP gimbal (EPMEC)

nism that provided two-axis re-orientation capabilityfor the electric propulsion thruster.

8.2.2. AOCS performanceThe majority of the problems in the AOCS occurred

in the Star Tracker at the beginning of the mission. Themost noticeable problems [2] were:

◦ Radiation damages, soft protons, severe solar storms,longer time cruising through the inner Van Allenbelts,

◦ Star Tracker high temperatures during cruise,◦ Rich star fields,◦ Star Tracker wrap-around, SEU/SEL (Micro Latch-

Up of CH#2),◦ Star Tracker time synchronization anomaly.

In general, the behaviour and pointing performance ofthe SMART-1 AOCS has been as expected throughoutthe mission. No hardware failures have occurred andonly one software error has been found and corrected.A few parameters such as thresholds were tuned in theinitial phase of the mission. For more detailed descrip-tions of the AOCS performance, see [12,13].

The majority of the problems in the AOCS occurredin the Star Tracker in the beginning of the mission and afew software modifications and parameter tunings wereaimed at compensating for these problems.

For the most noticeable problems of the Star Tracker,see [2]. The major part of these problems was caused bythe fact that the temperature of the Star Tracker cameraswas higher than expected. This in combination with anunexpectedly high radiation dose made the operation ofthe Star Tracker very difficult in the early phases.

Table 2 shows a cross relation between the anomaliesobserved and the cause of them. It is clear that noneof them were related to hardware problems and most


of them were caused by the environmental conditions.All of them were overcome via software patches.

8.2.3. AOCS operations experienceThe AOCS operations can be summarized as follows:

weekly preparation and uplink of attitude profiles, EPcommands, Sun vector commands and mode transitions,weekly monitoring and update of the gyro calibration,monthly ST performance monitoring, hydrazine book-keeping and monthly monitoring of the AOCS perfor-mance.

In the initial phase of the mission, the SMART-1 mis-sion control team/industry/project had to perform somespecial operations to tune the AOCS and cope with thehazardous space environment and new spacecraft char-acteristics: changes to some tuneable parameters in theAOCS and FDIR cores, permanent disabling of someAOCS FDIR functions related to the Star Tracker. Inaddition, push-broom operations for the AMIE payloadcamera required special spacecraft pointing to keep theStar Tracker CCD temperature below its operationallimits and finally procedures had to be developed forthe delta-V manoeuvres executed at the end of June andJuly 2006 to change the SMART-1 impact date and lo-cation on the Moon.

8.2.4. AOCS lessons learnt◦ The Star Tracker anomalies can for the most part not

be blamed on the actual unit: the CCD radiation dam-age was caused by the unexpectedly high radiationdose during the three months travel through the radi-ation belts. This was primarily caused by the severesolar activity during the fall of 2003. The only lessonlearnt could be to keep the temperature of the CCDas low as possible or to limit the exposition time inthe radiation belts.

◦ The problem with the temperatures outside opera-tional range could be overcome by improving designof the Star Tracker baffles or by active cooling.

◦ The Star Tracker processor overload could be reducedby installing one processor per camera as it is beingimplemented for the GOCE mission. This is probablynot necessary for the latest generation of the DTUStar Tracker since it is equipped with a more powerfulprocessing unit that is able to serve up to four cameraheads simultaneously.

◦ It has been difficult to analyse the Star Tracker be-haviour without insight in the Star Tracker softwarebut on the other hand the support from the manufac-turer (DTU) has been very efficient.

◦ The automatic power-toggling function of the StarTracker should be kept.

◦ The specific AOCS FDIR functions were rarely trig-gered, and never directly for the intended reason.

◦ Reconfiguration due to thermal anomaly was almostnever due to the fact that both units were at the sametemperature.

◦ More emphasis on data-handling: faulty data sent byunit should not trigger FDIR checks mend for ‘physi-cal anomalies’ (and sometimes cause a chain reactiontoo).

8.3. Radio frequency (RF) subsystem (TT&C)

The RF subsystem operates in S-band and supports:

◦ reception and decoding of software and pulsetelecommands;

◦ multiplexing and transmission of real-time and storedtelemetry at low rate (2 kbps) or high rate (66 kbps);

◦ ranging and Doppler measurements;◦ SAFE mode with autonomous toggling of the RF

switches once per hour to protect against single pointfailures.

The principal components are:

◦ 2 Omni-directional low gain antennas (LGA).◦ 1 Medium gain antenna (MGA).◦ 2 transponders including diplexers.◦ 3 co-axial switches.◦ 2 TMTC cards.◦ 2 oven controlled crystal oscillators (OCXO).

8.3.1. RF operationsEntry into the Safe mode activates an autonomous

RF switch toggling sequence to protect against singlepoint failures.

No RF single point failures occurred during the mis-sion but the autonomous reconfiguration of the subsys-tem was seen to work well after any entry into SAFE

mode. Use of LGA for high rate operations: for approx-imate the first 11 months of the mission the telemetrylink margins allowed SMART-1 to operate in high ratevia the omni-directional LGA antennas. This provideda higher degree of flexibility with regard to mission op-erations planning allowing telemetry stored in the on-board stores to be downloaded during every groundstation pass inside working hours.

It should be noted that long-term use of the LGA fortransmission of high rate telemetry was not a baselineoperational configuration (Fig. 12).

Eventually the increasing slant range associated withthe spiralling orbit to the Moon necessitated changes in


Fig. 12. Transponder–antenna connectivity.

the configuration of the RF subsystem as the ability torun high rate data via the low gain antennas degradedand was eventually been lost on day 2004.217 onwards(except via the 34 m antenna at New Norcia).

With the need to monitor the final EP thrusts moreclosely (day 2005.266 onwards) it became necessary toswitch to full low rate when the MGA was not availableat high slant range. Although this allowed continuousmonitoring of the spacecraft, subject to ground stationavailability, it resulted in frequent loss of Reed Solomonencoding due to the anomaly described in the nextsection.

From this point onwards the spacecraft was routinelyswitched between low and high rate telemetry rate as re-quired. Reed Solomon was set to passive at the groundstations resulting in a 2.8 dBm reduction in the link mar-gin, reception of bad telemetry frames at low elevationand occasionally during spacecraft slew manoeuvres.

8.3.2. RF performanceThe performance of the RF subsystem has been

broadly consistent with pre-launch modelling and bud-get calculations.

The main significant RF-related anomalies observedduring the mission were:

◦ De-synchronization of Reed Solomon encoding.◦ Minor degradation in performance of nominal trans-

mitter.◦ The transponder dual lock problem.

A small number of less significant anomalies have alsobeen observed and are also described in this section.

De-synchronization of Reed Solomon encoding: Afterlaunch a problem was detected with the Reed Solomonencoding on-board the spacecraft. It transpired that anyFPGA reset, CON warm/cold reset or even a teleme-try rate switch could cause the Reed Solomon encodingto lose synchronization and effectively became corrupt.Since corruption of Reed Solomon encoding could betriggered by routine changes of telemetry rate, opera-tions were conducted for most of the mission with ReedSolomon decoding set to passive at the ground station.

Additional filtering implemented within the MCSstopped the majority of this data reaching the rest ofthe telemetry processing chain. In general, operatingwithout Reed Solomon proved to be workable thoughthe potential remained for corrupt telemetry to reachthe MCS/DDS occasionally.

Degradation in output power of nominal transmitter:A degradation of about 1 dB in the output power oftransponder A was observed during the first nine monthsof the mission. This had negligible impact on operationsas the transponder was almost 1 dB above specificationat launch.

Dual lock problem: The problem of intermittent lockon the redundant telecommand receiver was experi-enced throughout the mission—this is the so-calledXMM problem.

In this situation a good lock was obtained on theprimary receiver and a marginal lock on the redundant


receiver connected to the antenna oriented away fromthe ground station. During telecommand uplink a com-mand might be routed via the redundant receiver andbecame corrupt.

As a result the dual lock problem could cause on-board command decoding failures, command retrans-missions and eventually a collapse of the commandingprotocol. If this happened in the middle of a missionplanning uplink then the recovery had to be a manualand time consuming process.

8.3.3. RF lesson learntThe overall performance of the RF subsystem was

good. The toggling mechanism after spacecraft separa-tion and transition to Safe mode worked as expected.

The main lessons learnt on SMART-1 are derived bythe anomalies detected;

◦ The Reed Solomon encoding design problem had se-rious consequences in the operations and in the costssince it required an expensive patch in the main con-trol system on ground.

◦ The hardware cross trap of the RF and bit lock indi-cations of the command link control word (CLCW)coming from the transponders was not implementedon SMART-1. This would have caused problem onground in case of a transponder failure. Future im-plementation should correct this hardware design.

◦ Finally, the dual lock problem is an old problem de-tected first during the XMM mission. This requires adefinitive solution for the future missions.

8.4. Thermal subsystem operations and performance

A traditional thermal concept with thermistor con-trolled heaters backed up by H/W thermostats on criticalequipment was used. Only passive thermal hardware asMLI, heat pipes, thermal washers, interface fillers andvarious coatings were used.

Thermal hardware:

◦ MLI: carbon filled Kapton was used.◦ Heat pipes: the heat pipes were screwed to the S/C

panels (+Y and −Y ).◦ Heaters: Kapton thermfoil with and without alu-

minium backing.◦ Thermostats: placed close to the thermistors.◦ Thermistors: a total of 58 thermistors were connected

to “TCRTU-A” and a similar set to “TCRTU-B”.

Thermal software:

◦ The thermal modes were: high, low, cell fail, EP ON.◦ Heater line control: all controlled by thermistors.

◦ If any of the thermistors used in the S/W thermostatfunction logic of a heater line has a temperature thatis lower than the low set point, then the power of thatheater line is switched ON.

◦ If all of the thermistors used in the S/W thermostatfunction logic of a heater line have a temperature thatis higher than the high set point, then the power ofthat heater line is switched OFF.

8.4.1. Thermal performanceThe performance of the thermal subsystem has been

very good during the entire mission even in extremesituations like i.e. low Moon orbits.

Solar array temperatures at end of mission: Aroundthe Moon, twice a year the Sun direction coincidedwith, or was close to, the orbital plane and the perilunewas not in the illuminated side of the Moon. Duringthis period the temperature of the arrays increased toa point, just before the Moon impact, where the fore-cast was above the qualification level. The analysis doneby ESOC/ESTEC/SSC concluded as a solution to turnthem away from the sun about 35◦. This special op-eration caused their temperatures to drop about 15 ◦C.The power generation reduction did not imply any newconstraint to the nominal operations.

Counter effect of long illumination periods of the−Z face of the S/C: There was a strong correlation be-tween sun on the −Z face, increase of the equipmentdeck temperatures and the high battery temperatures.This was noticeable when a special payload pointing re-quests were executed without a thermal mode of the S/Cchange. The analysis indicated that the primary causeof the high battery temperatures seemed to be long pe-riods of −Z face illumination, perhaps augmented byhigh temperatures of some units when the EP was off(thermal mode ‘High’).

Two corrective measures were implemented. Thefirst one was a change on the attitude in order to limitprolonged periods of −Z face solar illumination, andthe second was to lower the thermal settings of someunits of the equipment deck in thermal mode high (e.g.PCDU/BME). It was also tried with the hydrazine tankbut as the influence was negligible, it was decided tokeep the original settings.

8.5. Power subsystem

The main driver for the design of the power systemwas a requirement of EP, which required 1400 W ofregulated power. This resulted in the selection of a 50-Vregulated bus based on the S3R principle (sequencingswitching shunt regulator).


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The main components were:

◦ A two wings (2X3 panels) solar array generator.◦ A five identical battery management electronics sys-

tem (BME).◦ Five Li-Ion cells, with 135 W h capacity.◦ A power conditioning and distribution unit (PCDU):

40 low power solid state power controllers (LPSSPC),one high power solid state power controllers(HPSSPC).

◦ A main power controller (MPC).◦ One safety shunt (resistive dump).◦ 12 solar array shunts.◦ A PCDU power interface module (PPIM), which al-

lows to connect the EP-PPU directly to the main busin the event of HPSSPC failure.

The solar array generator including retention and de-ployment mechanism was provided by Dutch Spacebased on the FRED array developed for the ATV pro-gramme. Tecstar was the subcontractor for the photo-voltaic assembly (PVA) and solar array substrates. Thehigh efficiency TEC1 triple-junction cells solar cellswere of type GaInP2/GaAs/Ge.

The five BMEs and associated lithium batteries areused as part of a three regime control system that pro-duces a 50V bus regulated power bus in sunlight and ineclipse. During the sunlight regime the bus is regulatedby a reliable (triple-majority voted) 1 error amplifier lo-cated in the PCDU using a S3R solar array regulationsystem. During eclipse entry and exit, the BCR regimeis used to control the bus by reducing the load demand

on the power bus. In sunlight, the current from the 12solar array sections is fed into the solar array shuntsand the un-shunted current filtered by the main buscapacitor.

The solar array power from the array was slightlyabove 1850 W at the beginning of the mission.

8.5.1. Power subsystem performanceThe performance of the power subsystem has been

consistent with pre-launch power budget calculations.The power mode transition has been perfectly handledeclipse passages and during the EP thrust arcs.

From the entire mission, the main significant powerrelated observations were the degradation of the solararray (expected) and the drop of one of the solar arraysub-sections.

Fig. 13 shows the eclipse profile during the entiremission and the maximum eclipse duration (2 h 10 min)in April 2004.

The circle in Fig. 13 shows the maximum eclipseduration period and it was the only eclipse period whererequired special operations were performed.

The strategy followed was certainly to fly in an at-titude with −Z face of the S/C pointing to the Sun.This has had the beneficial effect of warming thebatteries before the long eclipses, increasing their ca-pacity and reducing the heater power consumption dur-ing the eclipse. In the longest eclipse seen so far onlyfour heaters were powered on. The power consumptionthrough the first ∼ 75 min of the eclipse stayed con-stant, after which some heaters started switching on.Throughout the eclipse only four heaters were used,


these being DCIXS, the small Xenon tank for the EPSand two heaters for the hydrazine system.

After reaching the final Moon orbit at the beginningof 2005, the eclipse seasons were very well defined andthe worst case of each one was when the Sun directioncoincided with, or was close to, the orbital plane andthe perilune was on the illuminated side of the Moon.During that time the S/C was in between the Sun andthe Moon for some time every orbit, making also anoticeable effect on the solar array temperature as it isexplained in the thermal section.

8.5.2. Power subsystem lesson learntThe solar array degradation expected after crossing

the radiation belts was about 18.5% (worst case) whilethe real one observed was just 8.5%. It was mainly dueto the long radiation influences experienced during thefirst part of mission crossing the Van Allen belts. Thedegradation performance of the triple junction solarpanels has been lower than expected.

A drop on the generation current of one of the solararray was observed since October 2005. It was causedby a miss performance of one of the solar array subsec-tions which was composed by 2 or 3 string of the totalof 98. It was assumed a possible lost slip ring in theBAPTAs or a failure in the solar array harness. From theavailable telemetry was not possible to define the reasonof this drop in the current generation. A more completeset of power operational telemetry would be useful tohave it, especially for degradation or failure analysis.

8.6. Data handling subsystem operations andperformance

The data handling subsystem contains cold redun-dancy, with autonomous failure detection isolation andrecovery (FDIR) software handling any single failures.The main controller runs on a 32 bit CPU ERC32 Sin-gle Chip. The remote terminal units (RTUs) are con-nected to a redundant commercial bus (CAN) to inter-face all units. The on-board software (OBSW) was de-signed with a high level of autonomy such that ground-command-sequence uplinks could be performed nomi-nally once every four days.

The controller (CON) is based around the AtmelTCS695E Single Chip ERC32 processor, with a 20 MHzsystem clock. The permanent storage memory area isprovided by 2 Mbyte EEPROM devices. A 4 Gbit Dy-namic RAM, EDAC protected, mass memory is used tostore on-board generated telemetry, temporary storageof uploaded software, for storage of control data, andfor on-board telecommand schedules.

8.6.1. Data handling performanceThe data handling system has performed nominal.

There were several double EDAC errors in the solidstate mass memory during the mission. The first threein 2003 could be related to the radiation effects insidethe Van Allen belts [1]. The following seven doubleEDAC errors occurred in Moon orbit and could not becorrelated with any external event or even other ESAmission.

No platform bus reconfiguration occurred during thefull mission. The payload bus was affected by a restric-tion in the OBSW that caused on two occasions a switchto the redundant bus. The switch to the redundant buswas not affecting the payload activities.

8.6.2. Data handling operations experience andlessons learnt

SMART-1 is using the standard ESA Packet Utilisa-tion Standard (PUS) Service for uplink and downlink.PUS Service could not be implemented for time devel-opment reasons impeding the development of on-boardcontrol procedures (OBCP) as in other ESA missions.

◦ Any on-board software patch should be permanent,especially in case of cold redundancy like SMART-1.

◦ Cold redundancy units should be patchable withoutinterrupting nominal operation in prime unit.

◦ On-board flow control priority could not be config-ured from ground; this had particular relevance inthe report packets generated for the on-board timetag queue (TTQ) and impeded their complete recep-tion on ground, obliging to request their dump 2 or3 times.

◦ PUS Service 1 used for the on-board confirmationof the commands on-board was also of low prior-ity and was frequently lost during payload/platformdumps. This had an important consequence in oper-ations since it was not possible to uplink the missionplan when a HK or payload dump was running, forc-ing the mutual exclusion of these operations.

◦ Another important subject for improvement is theconstraints on operations imposed by the fact that thehousekeeping data dumps could not be interruptedonce started.

◦ The commanding software of the payload unitsshould allow the use of grouping commands in akind of macro-command concept. This would havebeen extremely helpful during the Moon phase wherethe high number of science requests made a hugeincrease of the commands to uplink in every scienceplan.

◦ The events were grouped in long event messages thatwere sent every time any of them changed. This could


Fig. 14. Stores structure before the Moon phase.

not be changed during the mission and did its iden-tification very cumbersome.

8.6.3. Solid state mass memory configurationAfter the Moon capture, the configuration of the

solid state mass memory was changed to cope with theincrease of the science return data originated by thechange of the orbit date rate and also, to simplify thepayload downlink operations. The implemented ‘pay-load super store’ concept permitted to start and stop thepayload dump at any time leading to its full automationlater in the mission.

The new design based on just a read pointer and awrite pointer was implemented after the spacecraft ar-rival to the Moon operational orbit in March 2005. Itrequired a major patch on-board with the correspon-dent mapping of parameters on the ground (Figs. 14and 15).

8.7. Spacecraft data analysis, the mission utilitysupport tool (MUST)

The amount of anomalies at the early stages of themission and the need to quickly pass the information toexperts disperse around Europe put in evidence that theavailable tools were not appropriate.

This accelerated the development of a utility co-funded by the general studies programme of ESA.

MUST is a collection of tools that supports the exporta-tion, analysis and visualization of spacecraft telemetryand ancillary mission data via Internet (Fig. 16).

The system developed for SMART-1 demonstratedits cost-effective design and its versatility to be usedby other missions. It combines spacecraft telemetry andother ancillary data with web interfaces and alarmingsystem (subset also available to mobile phones andPDAs). Its implementation has been done securing theintegrity of the source data without affecting the perfor-mance of the MCS at ESOC.

8.8. Ground stations

The initial plan of having two passes per week lasting8 h was quickly abandoned in the early states of themission. The approach adopted to use any ESA stationbased on spare capacity. This was very beneficial for theoperations, the anomaly recoveries and the volume ofpayload data during the cruise and Moon phase. On theother side, the booking of ground station became oneof the most demanding activities for the flight controlteam.

SMART-1 ended up making use of many stations:Vil- 1, Vil-2 and TS-1 in Villafranca, Maspalomas(Canary Islands), Kourou (French Guyana), Perth andNew Norcia in Australia and Weilheim in Germany(via SLE).


Fig. 15. Stores structure during the Moon phase.

Fig. 16. MUST GUI and internet visualization of S/C TM.

8.9. SMART-1 ground operations automation system(GOAS)

The activities which could not be automated withthe mission planning system were automated with anautomation module.

The activities in question were:

◦ Routine download of platform housekeeping teleme-try and events.

◦ Intelligent download of payload telemetry.


Smart- 1 Operations Automation – 16/03/2006 ESOC (O.Camino, W.Heinen, R.Blake, J.Fortuno)

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◦ Download of gaps in platform housekeeping eventstelemetry prior to deletion of the on-board stores anddownload of gaps in payload telemetry.

◦ Real-time anomaly diagnosis based on out-of-limitsand telemetry values.

The above activities had been performed manually bymembers of the FCT according to existing flight con-trol procedures (FCPs). This required repetitive, man-ual effort on the part of the FCT and, in the case ofSMART-1 which operated without the traditional SPA-CON shift-based support, restricted these operations tonormal working hours (see [14]).

The GOAS started to be used operationally by the endof the mission in May 2006. Its architecture is depictedin Fig. 17.

9. Conclusions

The first of the ESA SMART missions, SMART-1 hasnot only achieved its primary objective of testing andvalidating operationally the use of electric propulsionfor interplanetary missions, but it has fulfilled all sec-ondary objectives in what concerns technology demon-stration and science.

The mission has permitted ESA gaining a valuableexpertise in navigation techniques for low thrust propul-

sion systems and experimenting with innovative cost-effective operations concepts for a fraction of a mediumscience mission (120 M¥).

The experience and records accumulated during theSmart-1 mission will allow ESA to undertake the nextgeneration of missions with confidence of success. Thelevel of operations automation achieved on both space-craft and ground has been remarkable and have set upnew benchmarks for the future.

SMART-1 has in addition awaked an unexpected levelof international interest, thanks to its achievements, andits potential usage for international cooperation, in par-ticular with cross support with DLR, serving as track-ing target t in spring 2006 with China CLTC and inJuly with the Indian ISRO for Moon exploration coop-eration and being used internationally for VLBI, deltaDOR and Ka band experiments.

References

[1] R. Furnell, et al., Smart-1 ERC32 In-Flight Performance &Lesson Learnt. First Annual ESA Workshop on Spacecraft DataSystems and Software, ESTEC 2005.

[2] M. Alonso, et al., Smart-1 Lunar Mission: Star TrackerOperations Experience. GNC Conference Loutraki GreeceOctober 2005.

[3] J. Schoenmaekers, Post-Launch Optimisation of The Smart-1Low-Thrust Trajectory to the Moon. Proceedings of the 18th


International Symposium on Space Flight Dynamics (ESA SP-548). Jointly organised by the German Space Operations Centerof DLR and the European Space Operations Centre of ESA.11-15 October 2004, Munich, Germany, p. 505.

[4] R. Rigger, et al., The Optimisation of Attitude Profiles forSmart-1: A Highly Constrained Problem (The Truth AboutAttitude). Proceedings of the 18th International Symposium onSpace Flight Dynamics (ESA SP-548). Jointly organised by theGerman Space Operations Center of DLR and the EuropeanSpace Operations Centre of ESA. 11-15 October 2004, Munich,Germany, p. 285

[5] M. Alonso, et al., Smart-1 Lunar Mission: OperationalExperience with its Automatic Attitude and Orbit ControlSubsystem and its Relation with Electric Propulsion System.GNC Conference Loutraki Greece October 2005.

[6] D. Gestal, et al., SMART-1 operations, IAC-05-B5.3/B5.5.[7] D. Milligan, et al., SMART-1 electrical propulsion operational

experience, IEPC-2005-245.[8] J. De Bruin, et al., Smart-1 lunar mission: from capture to

impact, IAC-06-A3.1.6.[9] C. Koppel, et al., The Smart-1 electrical propulsion subsystem

around the Moon: in flight experience, AIAA 2005–3671.

[10] D. Milligan, et al., Operationally enhanced electric propulsionperformance on electrically propelled spacecraft, IEPC-2005-246.

[11] P. Bodin, et al., Development, test and flight of the SMART-1 attitude and orbit control system, AIAA-2005-5991, AIAAGuidance, Navigation and Control Conference and Exhibit2005.

[12] P. Bodin, et al., Initial in-flight performance of the SMART-1 attitude and orbit control system, Proceedings of theInternational Lunar Conference 2003/ILEWG 5, AAS Scienceand Technology Series, vol. 108, AAS, San Diego, CA, 2004,pp. 317–328.

[13] P. Bodin, et al., The SMART-1 attitude and orbit controlsystem: flight results from the first mission phase, AIAA-2004-5244, AIAA Guidance, Navigation and Control Conference andExhibit 2004, AIAA Meeting Papers on Disc, vol. 9(15), AIAA,2004.

[14] W. Heinen, et al., Smart-1 ground operations automationspaceops, AIAA (2006) Ref 5710.

SMART1 operations experience and lessons learnt

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