Instrument Host Overview========================Data obtained from the Mars Express instruments were sent to groundvia the spacecraft on-board computer. As spacecraft to Earthcommunication does typically exclude instrument operations, all dataare relayed from the instrument to the spacecraft mass memory, thesolid state mass memory (SSMM). The data is downlinked to Earth viathe telemetry subsystem using ESA's antenna in New Norcia, Australia,and the Deep Space Network (DSN) antennas of NASA. The radio scienceexperiment requires data from the New Norcia and the DSN groundstation hardware.This catalogue file gives an overview of the spacecraft and theground stations used.For more detailed information see the spacecraft user manual,MEX-MMT-MA-1091.Instrument Host Overview - Spacecraft=====================================The spacecraft baseline design was the combination of missioncustomised configuration and mechanical / thermal architecture with aRosetta inherited avionics. The spacecraft design was driven bymission requirements, science return  and system concept. Aspacecraft articulation concept with body mounted instruments, fixedHigh Gain Antenna and 1 degree of freedom steerable Solar Arrays wasbaselined. The spacecraft design was based on a parallelipedic likeshape sizing  about 1.7 m length, 1.7 m width and 1.4 m height. Thesolar array was composed of two wings, providing a symmetricalconfiguration favourable to aerobraking  techniques and minimisingtorques and forces applied on the arrays and the drive mechanismsduring the Mars insertion manoeuvres performed with the main engine.So as to offer the adequate dry mass / propellant mass ratio andlarge mounting surfaces and volumes for the Orbiter instrumentsnecessary for Mars Express, the traditional cone/cylinder centralstructure has been found less efficient than a dedicated structuralconcept with only a Launch Vehicle Adapter connected to a stiffened'box' as now developed for light weight satellites. Within theoverall integrated design of the spacecraft, four main assemblies areplanned to simplify the development and integration process:(1) the Propulsion Module with the core structure,(2) the Y lateral walls, supporting the spacecraft avionics and the    solar arrays,(3) the Y/+X shear wall and the lower and upper floors, supporting    the payload  units. The +Zb face was nominally Nadir pointed    during science observation and Lander communication relay phases    around Mars, and supported Beagle 2 (released prior to Mars    capture) the Lander(s) relay antenna and ASPERA-3, and(4) the X lateral walls supporting the High Gain Antenna (-X) and the    instruments radiators (-X).Attitude and Orbit Control was achieved using a set of star sensors,gyros, accelerometers and reaction wheels. A bi-propellant reactioncontrol system was used for orbit and attitude manoeuvres by either a400 N main engine or banks of 10N thrusters. The Data Handling isbased on packet telemetry and telecommand. The Electrical Powergeneration was performed by solar arrays, the power storage by aLithium-Ion battery. A standard 28 V regulated main bus is offered tothe payload instruments. The RF Communications function transmittedX Band telemetry 8 hours per day via a High Gain Antenna at ratesbetween about 19 and 230 kbps depending of the Mars to EarthDistance. A variable telecommand rate of 7.81 to 2000 bps wasforeseen during up to 8 hour per day.Spacecraft Coordinate System----------------------------The origin of the spacecraft Reference Frame, named Oa, was locatedat the separation plane between the spacecraft and the adapter, atthe centre of the interface diameter of 937 mm.-  The Xa axis was contained in the spacecraft/launch vehicle   separation plane, and oriented toward the High Gain Antenna side   of the spacecraft.-  The Za axis was coincident with the launcher X1-axis. It   represents the SC line of sight toward Mars during science   operation, and the ejection direction for the Beagle 2 probe.-  The Ya axis was contained in the SC/LV separation plane, and   oriented so as to complete the right handed co-ordinate system. It   is therefore parallel to the solar array plane and positively   oriented opposite to Marsis antenna support wall.The (Ob, Xb, Yb, Zb) Reference Frame is structure related, and is notused at S/C or operations level.Spacecraft Structure and Interface With Payload Units-----------------------------------------------------The selected SC structure limits the number of complex elements tothe bare minimum. Indeed, the only cylindrical part of largedimensions was the Launch Vehicle Adapter ring, the rest of thestructural items being principally flat, standard panels withaluminium skins and aluminium honeycomb. The structure was composedof:1) a Core Structure, built up from : - One Launch Vehicle Adapter ring machined from a solid aluminium   cylinder of approx. diameter 940 mm, 200 mm height with a   thickness of 3.5 mm. This LVA was the main load path transfer from   the Spacecraft body to the launch vehicle interface. - Two Tank Beams supporting the lower tank bosses, and embedded in   the LVA ring, - Two Upper Tank Floors, supporting the tanks upper bosses, - One Lower Floor. - Two X Shear Walls, - One Shear Walls in the Y direction2) an Outer Structure, built up from : - a +Z Top Floor, - two +Y and -Y Sidewalls, - two +X and -X Lateral Closure Panels (split to allow separate   access into each quadrant), - various dedicated equipment support panels (PFS, Omega and   pressurant tank) - miscellaneous brackets (e.g. to support sensors, antennas,   propulsion items, instruments).All these elements were made of Aluminium Alloy, either from forgings(LVA ring, tank beams and main brackets) or from honeycomb sandwichpanels. The panels were made of honeycomb (generally type 1/8-5056-0.001P) of 10 to 20 mm thickness, bonded to Aluminium facesheets ofthickness varying between 0.2 to 0.3 mm and up to 0.5mm additionaldoubler for local reinforcements.In general, the payload units were accommodated following their mainneeds. The payloads needing a stringent thermal control and/orpointing performances (HRSC, OMEGA, PFS, SPICAM) were gathered on, orclose to, the +X shear wall, inside the spacecraft and close to theAOCS reference (namely the Inertial Measurement Package and the StarSensors). In order to meet the PFS scanner to PFS sensor co-alignmentwithout disturbance caused by dismounting, these PFS units wereinstalled on a stiff, removable mounting assembly which can beintegrated as a single unit on the spacecraft. To expediteinstallation of the large Omega-SA, this instrument was installed viaedge-mounted inserts in the Y-shear wall and a dedicated Omegasupport panel.The payloads requiring a large field of view and not necessitatingstringent thermal control were located externally on the Top andBottom Floors (ASPERA) or lower edge of the +Y sidewall (MARSIS).None of the payload units required isostatic mountings: they wererigidly fixed to the spacecraft structure utilising standard space-industry inserts and screws.Some of the payload units were of significant mass and thereforerequire the implementation of large, face-to-face inserts that arebonded inside the sandwich panels at the time of panel moulding.Beagle 2 was accommodated on the top floor of the S/C, in order tominimise dynamic disturbance (centre of mass transfer in the (X, Y)plane) and then maximise the reliability of the Mars orbit insertionmanoeuvre. The remaining Beagle 2 hardware after probe ejection isconstrained within 50 mm height and is thermally insulated tominimise straylight and thermal distortion disturbances respectively.Thermal Control---------------The spacecraft thermal control was in charge of maintaining allspacecraft equipment within their allowed temperature ranges duringall mission phases. The equipments fall into two categories: - the collectively controlled units, for which the heat rejection   and heating capabilities (design and accommodation) are provided   by the spacecraft thermal control, - the individually controlled units, self provided with their own   thermal control features (coatings selection, heaters,   insulators), for which the spacecraft thermal design controls the   thermal interfaces within the required ranges.A passive thermal control design was implemented for the Mars Expressspacecraft; it was supplemented with an electrical heating system.The heat rejection toward space was performed using radiators mainlyon the +/-Y panels for the platform internal units and the +X panelfor the payload equipments.These sides of the spacecraft are the most favourable areas, beingmost of the time protected from the direct sun inputs (always for the+X side). The Mars planet flux are imposed by the spacecraft orbitand attitude and mainly significant during the pericentre phase inoperation. The rest of the spacecraft is insulated with Multi LayerInsulation blankets to minimise the heat exchange and the temperaturefluctuations.The spacecraft external units (Platform and Payload units) werethermally decoupled from the spacecraft and provided with theirindividual radiator when needed. The electrical heater system allowedto raise the temperature of the unit above their minimum allowedlimits, with temperature regulation functions provided either bymechanical device or by the onboard software. Most of the spacecraftunits were collectively controlled inside defined thermal enclosuresin which the heat balances were controlled by proper sizing of heatrejecting radiators and heating power implementation. It allowed tomaintain the unit temperatures to acceptable levels. The heattransfer from the units to the radiators was performed by conductionwhen unit baseplates were attached to the radiator honeycomb panelsand by radiation. In that case units and panels had a black finish tomaximise heat transfer inside the thermal enclosures.For more demanding units like the HRSC and OMEGA cameras, and the PFSspectrometer, featuring their own thermal control, specialprecautions were taken by individual trimming of their conductive andradiative isolation. The HRSC camera required a temperature controlwithin a narrow temperature range):it was provided with a thermal strap connecting it to a dedicatedradiator tuned to limit the temperature excursion in operation withina 10 degree C temperature range.OMEGA and PFS are provided with dedicated radiators, implemented onthe +X side of the Spacecraft. Whatever the Sun / Earth / Mars /Spacecraft geometry, the +X side of the Spacecraft was oriented awayfrom Sun over the complete Martian orbit, both during Nadir pointedscience phase and Earth pointed communication phase.This allowed to provide the camera and the spectrometer with athermal interface at temperature lower than 175K and 190Krespectively during the Planet observation. The connection to theradiators were performed by thermal straps, the radiators beingthemselves decoupled from the rest of the spacecraft using thermalblankets and insulating stand-offs.Payload external units like MARSIS and MELACOM antennas, ASPERA-3units, were individually controlled units. They required large fieldof view and thus were directly affected by the external environmentand they had to withstand larger temperature ranges than the standardunits. They are as far as possible insulated from the spacecraft.Their coatings were selected and trimmed to suit. The spacecraftinterface temperature had a very limited influence on their thermalbehaviour.The propulsion equipments that were mounted internally were ingeneral isolated with MLI, and provided with their own thermalcontrol heaters: tanks, fluid lines, valves, pressure sensors. Themain engine and the thrusters had their thermal coupling with thespacecraft tailored to meet their thermal requirement whilepreserving the spacecraft thermal behaviour. They were provided withindividual electrical heaters sized to maintain these external unitswithin the acceptable temperature range accounting for wide change inradiative environment.The High Gain Antenna was using a passive thermal control: aKapton/Germanium sunshield was covering the whole antenna on itsfront side, while a light weight MLI is used on the rear side of thereflector.MECHANISMS----------The implementation of mechanisms into the spacecraft configurationhad been kept to the minimum. The mechanisms employed are thoseassociated with- Reaction Wheel Assembly (RWA),- Solar Array Drive Mechanism (SADM),- Solar Array Hold-Down and Release Mechanism (HDRM),- Solar Array Deployment System,- Beagle-2 Spin-Up and Ejection Mechanism (SUEM) and the- MARSIS antennas deployment mechanism.REACTION WHEEL ASSEMBLYThe Attitude and Orbit Control System (AOCS) of the spacecraftrequired implementation of four reaction wheels, used with a threeout of four redundancy. They were of ball bearing momentum / reactionwheel type, for clock-wise and counter-clockwise operation, with thewheel mass suspended by two angular contact ball bearings paired bysolid preloading. The main functions of the RWA was to ensure correctorientation of the spacecraft in fine pointing modes, and to ensurespacecraft manoeuvrability (e.g. at transition between Mars orbitobservation and communication phases), with minimum propellantconsumption (the only related consumption lied with wheel momentumoff-loading that had to be performed at regular intervals, typicallyevery 2 days.)SOLAR ARRAY DRIVE MECHANISMThere were two SADM used on the spacecraft, one for each Solar Arraywing. The SADM were mounted on each side of the spacecraft, and wereindependently controlled by the AOCS Processor Module. The mainfunctions of the SADM was to support the solar array wing throughoutthe mission, to provide the electrical power and signal interfaces tothe spacecraft and to orient the solar array wing towards the Sun byrotation about the Ys axis. The SADM was composed of a motor andgearbox assembly, ensuring the orientation of the solar array byrotation, a shaft and bearing assembly ensuring mechanical connectionand pointing accuracy, a twist capsule unit transferring electricalpower to the spacecraft. Those elements were mounted on a baseplatewhich was attached to the spacecraft sidewall.SOLAR ARRAY HOLD-DOWN AND RELEASE MECHANISMEach wing of the solar array was attached on the spacecraft sidewall,in launch configuration, by Hold Down and Release Mechanisms (HDRM).Each HDRM consisted in a set of hold down bushings, attached to thestructure of each panel which were held together via a stainlesssteel hold down pin of 3.5 mm diameter on a hold-down baseplate fixedon the spacecraft sidewall. The HRDM also incorporated a pair of pyroinitiators, which were actuated after spacecraft separation from thelauncher under control of the Data Handling Processor Module. Themain functions of the HDRM was therefore to maintain safely stowedeach solar array wing and to ensure their release for proper solararray power generation.SOLAR ARRAY DEPLOYMENT SYSTEMEach yoke and wing of the solar array was fitted with a deploymentmechanism that ensured proper deployment and latching of the solararray after release of the HDRM. The deployment mechanism consistedin a set of spring energy driven hinges mounted by pair between eachsolar array panel, between the first panel and the yoke, and betweenthe yoke and the SADM. Each hingeline was then linked to the othersby a set of pulley and cables, that ensured a synchronised deploymentof the wing.The torque margin of the Solar Array deployment system varied between7.5 (at beginning of deployment) and 2.6 (at end of deployment).BEAGLE 2 SPIN-UP AND EJECTION MECHANISMBeagle 2 formed an integrated experiment, composed of a lander(featuring investigation experiments) encapsulated in a Entry,Descent and Landing System (EDLS). Those items were composing theprobe, which interfaced to the orbiter top floor through the Spin-Upand Ejector Mechanism.MARSIS ANTENNAS DEPLOYMENT MECHANISMSThe baseline configuration for MARSIS deployment mechanism haddeparted from the Cassini (STEM) concept, i.e. a tubular antenna madeof 2 semi-circular formed strips made of Copper-Beryllium.The selected design was the ASTRO one, consisting of a boom made of aGFRP tube pierced with 2 diametrically opposed diamond shaped holesat the selected distance to provide folding capability. The antennaboom contained two wire elements forming the active radioelectricalpart of the antenna, and was folded at each hollowed hinge and heldflattened in specific containers. When release was initiated, thecontainer was opened through pyro devices, and the boom was selfdeploying thanks to its intrinsic energy which had been stored duringthe folding/flattening process necessary to meet the launch volumeconstraints.ATTITUDE AND ORBIT CONTROL SYSTEM---------------------------------AOCS BASIC CONCEPTSDue to the selection of a fixed high Gain antenna (HGA), and to thepropulsion configuration including a Main Engine, the Mars Expressmission required a high level of attitude manoeuvrability for thespacecraft. Attitude manoeuvres were performed:- Between the observation phase and the Earth communication phase, or  to reach specific attitudes necessary for science observations (in  particular SPICAM).- Before and after the Lander ejection, before and after each  trajectory correction manoeuvre, performed either with the Main  Engine or with the 10N thrusters.- To optimise the Wheel Off-Loading, through the selection of an  adapted attitude for this operation.All the attitude manoeuvres of operational phase were defined onground, using a polynomial description of the Quaternion to befollowed by the Spacecraft. The attitude estimation was based on StarTracker and gyros, ensuring the availability of the measurements inalmost any attitudes. Some constraints had however to be fulfilled,the Star Tracker being unable to provide attitude data, when the sunor the planet are close to or inside its Field of view. Reactionwheels were used for almost all the attitude manoeuvres, providing agreat flexibility to the Spacecraft and reducing the fuelconsumption. The angular momentum of the wheels had however to bemanaged carefully from ground.STAR TRACKER (STR)The Star Tracker (STR) was the main optical sensor of the AOCS, usedat the end of the attitude acquisition to acquire the final 3-axespointing, and during almost all the nominal operations of themission. A medium Field Of View (16.4 deg circular) and a sensitivityto Magnitude 5.5 were used to provide  a 3-axes attitude measurementwith at least 3 stars permanently present in the FOV.The STR included a star pattern recognition function and can performautonomously the attitude acquisition. The Mars Express Star Trackerwas produced by Officine Galileo, and is similar to the Rosetta one,except at S/W level. 2 Star Trackers were implemented on the minus Xaface of the Spacecraft, with an angle of 30 degree between theiroptical axes.INERTIAL MEASUREMENT UNITS (IMU)Two Inertial Measurement Units (IMU) were used by the AOCS, each IMUincluding a set of 3 gyros and 3 accelerometers aligned along 3orthogonal axes. The AOCS control used either the 3 gyros of the sameIMU (reference solution at the beginning of life) or any combinationof 3 gyros among the 6 provided by both IMUs. For the accelerometers,only a full set of accelerometers of one single IMU was used, due tothe lower criticality of the accelerometer function, and to theavailability onboard of an alternative method for the delta Vmeasurement (pulse counting). The Gyros were useful during theattitude acquisition phase for the rate control, during theobservation phase to ensure the required pointing performances andduring the trajectory corrections, for the control robustness andfailure detection. A non mechanical technology was selected to avoidthe mechanical sources of failure in flight. The Accelerometers wereessential during the main trajectory corrections such as theinsertion manoeuvre to improve the accuracy of the delta V. The IMUof Mars Express is identical to the Rosetta unit. Only the number ofunits and the onboard management of the configuration was different.SUN ACQUISITION SENSORS (SAS)Two redunded Sun Acquisition Sensors (SAS) were implemented on theSpacecraft central body and are used for the pointing of the SunAcquisition Mode (SAM) during the attitude acquisition orreacquisition in case of failure. The SAS are identical to Rosettaunits, but provided with customised baffles.REACTION WHEEL ASSEMBLY (RWA)The Reaction Wheel Assembly (RWA) included 4 Reaction Wheels (RW)implemented on a skewed configuration. This configuration enabled toperform most of the nominal operations of the mission with a 3 RWLconfiguration among 4. During some critical phase during which thetransition to the SAM had to be avoided (before lander ejection andbefore Mars Insertion Manoeuvre), a 4 wheels configuration was beused, under ground request. The Reaction wheels provided the AOCScontrol torques during all the phases of the mission except thetrajectory corrections, the attitude acquisition and back up modes.PROPULSION CONFIGURATIONThe Propulsion configuration included a Main Engine (414 N) which wasused to perform all the major trajectory changes, and 10 N thrustersused for the attitude control and also to produce the thrust duringthe small trajectory corrections. The 10 N thrusters configurationwas optimised to perform all the attitude control functions with only4 redunded thrusters, each of them being implemented near a corner ofthe -Z face of the spacecraft.SOLAR ARRAY DRIVE MECHANISM2 redunded Solar Array Drive Mechanisms (SADM) were implemented onthe Y+ and Y-walls of the spacecraft to control the orientation ofthe Solar Arrays. The SADM was only used for large angle orientationof the wings, the selected flight orientation during the observationphase near pericentre requiring no SADM actuation, once theobservation attitude was reached. The SADM used a stepper motor, agear, and a twist capsule technology. The SADM motion is defined inthe range +/-180 deg (minus margins). The SADM is identical to theRosetta unit, except for the speed levels which are specific to MarsExpress.AOCS HARDWARE ARCHITECTUREAOCS unit Nb   Technology / characteristics   Heritage       Supplier------------   ----------------------------   ----------     --------Star Tracker   2 CCD detector. 16.4deg        Rosetta unit.  Officine               circular FOV/ Magnitude 5.5                   GalileoGyro/accelero   2 Ring Laser Gyros (RLG).     Rosetta unit   Honeywell                3 gyros/3 acceleros per                unit.Sun Acquisition 2 Solar cells mounted on      Rosetta/SOHO   TPD-TNOSensor (SAS)      a pyramidReaction Wheel  4 Ball bearing Momentum/      Telecom. Sat.  Teldix                Reaction wheels.              Unit                 12 Nms/0.075 NmSADM            2 Stepper motor with gear.    Rosetta unit   Kongsberg                Twist capsuleAOCS GENERIC FUNCTIONS----------------------The AOCS modes used generic functions for the guidance, the attitudeestimation and the actuators management. The role of the guidance wasto provide onboard the reference attitude to be followed at each timeof the mission by the attitude control. It concerned of course theorientation of the Spacecraft but also the Solar Array position. Theanalysis of the mission needs showed that 4 types of guidance arenecessary along Mars Express mission:- Pointing of the High Gain Antenna (HGA) towards the Earth, and the  Solar Array cells towards the Sun. This kind of guidance was used  during the cruise phase and for communications during the  scientific mission phase, these two cases corresponding to the AOCS  Normal Mode, pointing on ephemeredes (NM/ GSEP phase).  The information necessary to the guidance concerned the Spacecraft  to Earth and the Spacecraft to Sun directions. They were contained  in the ephemeris definition.- This type of guidance was also used in a different way for the  Earth acquisition (SHM : Safe/Hold Mode), in order to perform the  autonomous orientation of the spacecraft towards the Earth. The  ephemeris data were then used to perform large angle slew  manoeuvres with thruster control.- Attitude profiles : this type of guidance was necessary during the  observation phase for the Nadir pointing or to follow more specific  profiles. This function was ensured by an onboard profile  description based on Chebychev polynomial, the parameters being  uploaded from ground. This capability enabled also to ensure the  attitude slew manoeuvres.- Fixed inertial pointing (fixed quaternion) : This type of guidance  was used for specific phases of the mission, during Orbit Control  Mode, Thruster Transition Mode or during the scientific mission  phase for SPICAM specific needs (in NM/FPIP and NM/WDP).Three generic functions had been defined for this purpose at softwarelevel :- the Ground commanded guidance,- the Onboard Ephemeris propagation,- the Autonomous Attitude Guidance Function, this latter function  generating the guidance information necessary either for the fixed  Earth pointing or for the Earth acquisition in SHM.GYRO-STELLAR ESTIMATION FUNCTIONThe gyro-stellar estimation function was common to many AOCS modes :It was initialised during the Sun Acquisition Mode (SAM) to preparethe following Earth acquisition operation (SHM: Safe/Hold Mode). Itprovided accurate attitude estimation during the Normal Mode ofcourse but also in the Orbit Control Mode (OCM) and ThrusterTransition Mode( TTM) for instance. The gyro-stellar estimatorprocessed gyro and star tracker (STR) measurements to provide anaccurate estimate of the spacecraft attitude. It was based on aKalman filter with constant covariance that allowed mixingmeasurements at different rates (8 Hz for the gyros and 2 Hz for theSTR). The constant covariance reduces the computer load whileensuring good performances. The estimated attitude was a quaternionrepresenting the spacecraft attitude in the J2000 inertial frame.The gyro-stellar estimator also estimated the gyros drifts to limitthe attitude errors in case of STR measurement absence due, forinstance, to a temporarily STR occultation. A specific management ofthe drift estimates was proposed for Mars Express, taking intoaccount the specific conditions of the scientific mission phase(existence of rates due to varying profiles, and potentialoccultation). The gyro-stellar estimator implemented a coherency testbetween the gyro and STR measurements in order to detect failuresthat could not be detected at equipment level.REACTION-WHEEL OFF-LOADING FUNCTIONThe wheel Off-Loading function enabled to manage the angular momentumof the wheels to a target value, through thruster actuations. Thisfunction was completely autonomous during the last phase of the Earthacquisition sequence (SHM/EPP:Earth Pointing Phase). During thenominal operations around Mars, it was preferable to command thewheel Off-Loading from the ground, the date being optimised takinginto account the mission constraints. The Off-Loading functionmanaged simultaneously all the wheels. It included several sequencesof thruster pulses until angular momentum of each wheel was close tothe target value. This sequence was defined by a feed forward 3-axeswheel torque command combined with a thruster pulse.The sequence ended with a tranquillisation phase controlled by thewheels, in order to damp the dynamic excitation generated by theactuation of thrusters and wheels.REACTION WHEEL MANAGEMENT FUNCTIONThis function was active in all the modes controlled through wheeltorques (Normal Mode and Safe/Hold Mode at the end of the attitudeacquisition sequence), but also when the wheels were kept to aconstant speed through a specific control loop but not used in theAOCS control, as in Orbit Control Mode, Thruster Transition Mode orBraking Mode. Six states of the wheel configuration are possible withthis function depending on the control of the wheels in torques (t)or in speed (s). For instance, the nominal operation in Normal Mode,uses 3 wheels in torques (3t), but could sometime require a fourthwheel if a hot redundancy is useful (4t). During trajectorycorrections the configuration included 3 wheels controlled in speed(3s). Intermediate states are necessary between these basicconfigurations in order to spin the wheels for instance (3t + 1s).This function was also in charge of the generation of wheel torquecommands in wheel frame, and of the friction torque estimationnecessary for compensation and for the failure detection. Itinterfaced also with the Wheel Off-Loading function.THRUSTER MODULATOR AND SELECTION FUNCTIONThe selected amplitude modulator and on-time summation algorithmswere re- used from Rosetta and adapted to match more efficiently theMars Express needs taking into account the specific thrustersconfiguration.The modulator had only one working phase where the four thrusters canbe used:- to produce a force along the satellite Z axis direction- to control the 3-axes satellite attitude (three torques are  commanded to the modulator).The modulator working frequency was 8Hz. At each step, the modulationtype used (ON-modulation or OFF-modulation) was automaticallyselected so as to maximise the available torque capacity for attitudecontrol. In the case the torque capacity was insufficient withrespect to the commanded control torque, priority is given to thecontrol and the commanded force ratio is automatically modified torecover the required torque capacity. Moreover in order to limit theactuation delay, the attitude control torque was always produced atthe beginning of the actuation period.To limit the number of thrusters ON/OFF or to tune the control limitcycle amplitude when using thrusters, the modulator output period hadto be changed to any period multiple of 125 ms.PROPULSION ARCHITECTURE DESCRIPTION-----------------------------------A bi-propellant system based on a telecommunication spacecraftheritage was adopted for the baseline. A set of isolation pyro valvesand latch valves had been added to ensure safe operations duringLaunch and Cruise, and for a re- liable acquisition of the Mars orbitfor science mission.At launch, the pressurant assembly (high and low pressure sections)were all isolated from the propellant tanks by normally closedpyrotechnic valves PVNC1 to PVNC6, by non return valves NRV1 to NRV4.The propellant tanks are pressurised to 4 bar. Similarly, thepropellant was isolated from the Reaction Control Thrusters and MainEngine assembly by normally closed pyrotechnic valves PVNC7 to PVNC14and thruster/main engine Flow Control Valves (FCV).Following separation, the pyro valves protecting the pressurantassembly were fired to pressurise the system to its operatingpressure of 17 bar. Then the latch valves were closed, isolating thenon return valves from propellant. A pressure transducer (PT2)located at the regulator outlet could monitor pressure build up atthe NRV location due to regulator leakage. When necessary the latchvalves were opened and the pressure relieved into the propellanttanks. It was assumed that a pressure of up to 20.5 bar could be thecriterion to initiate an open/ close cycle of the latch valves bytelecommand.The 20.5 bar pressure was an initial suggestion which needed to beconfirmed. It may affect component qualification issues because itexceeds existing MEOP values for the components in the section. Shortduration opening times for the latch valves minimised propellantvapour migration and it was essential for both oxidiser side and fuelside latch valves to open simultaneously to limit vapour migration.The system operates in this pressure regulated mode, using the 10 NReaction Control Thrusters only, during the period of the transferorbit to Mars.A few days before Mars orbit insertion, the 400 N Main Engine wasprimed and then calibrated by specific blank manoeuvres, combinedwith re-targeting of the S/C after Beagle 2 probe ejection. Thisensured that the Main Engine could be used safely for the Mars orbitinsertion and acquisition of the operational orbit. Should a MainEngine failure be detected at this stage, a back-up scheme, using theReaction Control Thrusters would have been implemented to reach atleast a degraded orbit around Mars.After attaining the operational orbit, the pressurant and Main Engineassemblies were re-isolated by firing all the normally openpyrotechnic valves and closing the latch valves. The remainder of themission was per- formed in blow down mode, using only the 10 NReaction Control Thrusters. The number of Reaction Control Thrusterhad been limited to 8 (4 nominal, 4 redundant), located at the bottom(-Z) side of the spacecraft to provide thrust principally along Zb tocompensate for Main Engine thrust imbalance caused by Main Enginealignment and Spacecraft Centre Of Mass (CoM) uncertainties. Adequatetilting of the Reaction Control Thrusters is implemented so as toprovide the capability for torque around each main axis of thespacecraft.In order to maximise flexibility and adaptability to failure cases,each Reaction Control Thruster was fitted with a Thruster Latch Valve(TLV) upstream from the thruster Flow Control Valves, permittingindividual switch over from prime to redundant for each ReactionControl Thruster. It had to be noted that this two-tank configurationwas compatible with a horizontal handling of the spacecraft asrequired by Soyuz launch campaign, on the proviso that the tanks werefilled at least up to 62% of their maximum capacity. Thecompatibility of this fill fraction wrt S/C global dynamic behaviourwas under investigation to avoid fluid/structural modes coupling.RF COMMUNICATIONS-----------------OVERVIEWThe communications with the Earth could be performed either in S-Bandor X-Band in accordance with ESA Standards. Two Low Gain Antennas(LGA) allow omni-directional emission and reception in S-Band, whilea dual band 1.65 m High Gain Antenna (HGA) allows high rate TMemission in S-Band and X-Band including TC reception in S-Band and X-Band. Demodulation of the up-link signal was performed by the DualBand Transponder before routing the resulting bit flow to the DataHandling. The stored TM within the SSMM is modulated in either SBandor X-Band within the Dual Band Transponder, which also performed S-Band signal amplification with 5 W. X-Band signal amplification isperformed using a 65 W Travelling Wave Tube Amplifier.UPLINKThe communication from the ground station(s) to the spacecraft wasperformed in S-Band or X-Band. Two Low Gain S- Band Antennas (LGA)were accommodated, one on the upper Z-panel, aside of the High GainAntenna and the other one on the bottom of the spacecraft, thusallowing a quasi omnidirectional coverage. The LGA was used mainlyduring Launch and Early Operation Phase (LEOP), critical phases andfor emergency situations. A narrow-beam dual-band high-gain antennawas used for all nominal mission operations for the uplink in X-Band,like the Cruise Phase or when orbiting around Mars.The RF uplink signal, which was modulated with packetisedtelecommands as NRZ/PSK/PM data, was routed towards a diplexer,performing frequency discrimination, and then to the Dual BandTransponder input. The transponder performed carrier acquisition anddemodulation, and transmitted the extracted signal to the DataHandling for further processing.The frequencies for the uplinks are:- 2114.676 MHZ (DSN 18) for S-Band,- 7166.936 MHZ (DSN 18) for X-Band.The following telecommand bit rates are handled by the Mars ExpressSpacecraft as provided by the CDMU design: 7.8125 bps and 15.625 bps,250 bps, 500 bps, 1000 bps and 2000 bps. These possible bit rates areselectable by Memory Load Command (MLC). As a baseline, the lowestbit rates was used in case of emergency via the Low Gain Antennas inS-Band, while the highest ones were used operationally through theHigh Gain Antenna in XBand.DOWNLINKA high data downlink capability was required, considering the largedata volume generated by the instruments. Nevertheless, downlinkcapacity was limited by the large spacecraft to Earth distance. Thedownlink of the telemetry data to the ground stations were performedin either S or X-Band.The frequencies for the downlink were:- 2296.482 MHZ (DSN 18) for S-Band,- 8420.432 MHZ (DSN 18) for X-Band.Downlink was performed at a commandable, variable bit rate. The CDMUdesign allowed to generate a telemetry flow at any bit ratecorresponding to a power of two multiplied by 32/n and lower than262.144 bps, where n is equal to 2, 3, 5 or 7. The possible bit rateswere selected via Memory Load Command (MLC) and vary from 7.8 bps asa minimum and can be up to 230 kbps. The bit rate to which referencewas made was the bit rate following Reed-Solomon encoding, but priorto convolutional encoding, if any. Due to hardware limitations,convolutional encoding was only performed for bit rates lower than65536 bps. Above this value, only Reed-Solomon encoding wasperformed.As a baseline, the lowest bit rates were used in case of emergencyonly using the Low Gain Antennas, whilst the highest ones were usedoperationally through the High Gain Antenna in X Band. The variablebit rate signal was transmitted to the Dual Band Transponder as SP-L/PSK for bit rates lower than 65536 bps and as SP-L (no subcarrier)for higher bit rates. This signal was phase- modulated in either SBand or X Band by the Dual Band Transponder, and added to the MPTSranging signal if it had been detected on the uplink.DATA HANDLING ARCHITECTURE--------------------------The Data Management System (DMS) was in charge of telecommanddistribution to the whole spacecraft, of telemetry data collectionfrom the spacecraft sub- systems and payloads and data formatting,and of the overall supervision of spacecraft and payload functionsand health.The DMS was based on a standard OBDH bus architecture enhanced byhigh rate IEEE 1355 serial data link between the CDMU (Control andData Management Units) processors and the SSMM and STR. The OBDH buswas the data route for platform and payloads data acquisition andcommands distribution via the RTU. The DMS included 4 identicalProcessor Modules (PM, 1 to 4) located in the 2 CDMU.  Two processormodules were dedicated to the DMS (PM2 and PM3), and two to theAOCS(PM1 and PM4). The PM selected for the DMS function acted as thebus master. It was in charge of Platform subsystem management(Communications, Power, Thermal, Payloads). The PM selected as theAOCS computer was in charge of all sensors, actuators and Solar ArrayDrive Electronics (SADE).TC-decoder and Transfer Frame Generator (TFG) were included in eachCDMU. The Solid State Mass Memory (SSMM) was used for data storageincluding 12 Gbits of memory at BOL. It was coupled to the two DMSprocessors, the TFG, OMEGA, HRSC and MELACOM instruments. It storesscience and global housekeeping telemetry packets.OVERVIEWThe Data Handling architecture was organised around the two CDMU.They were in charge of controlling ground command reception andexecution, on-board housekeeping and science data telemetry storageand formatting them for transmission. The on-board data management,controlled processing and execution of on-board control proceduresbelongs to their tasks as well. Each CDMU featured two MA3-1750Processor Modules, each of them being able to process either DataManagement or AOCS software.A built-in failure operational Reconfiguration Module (RM) ensuredsystem level FDIR and reconfigured the CDMU as necessary. Datatransfer with other Data Handling units were ensured using standardlinks such as a redunded OBDH data bus or IEEE-1355 serial links. TwoInterface Units were performing inter- face adaptation between thoselinks and other spacecraft units. The AOCS Interface Unit (AIU) wasdedicated to AOCS equipment, while the RTU interfaces with theremainders, including the Instruments. A file-organised 12 Gbits SSMMwas implemented to store the Housekeeping and the Science Data. Italso collected directly Science Data from the three high rate PayloadInstruments.SSMM SOFTWARE-------------The Solid State Mass Memory (SSMM) consists of 2 processor systems:- The Memory System Supervisor (MSS), dedicated to the communication  with the DMS MMS.- The File and Packet Controller (FPC), dedicated to the file  management on the memory modules and to the data exchange with the  instruments and the TFG.The SSMM software runs on the micro-processor based MSS and themicro- controller located in the FPC. The main part of the SSMM-SW isprogrammed in C language. Parts of the start-up function areprogrammed in Assembler. The SSMM software consists in two parts:- The Initialisation software covering the Init Mode and running in  the MSS. It was executed in MSS PROM after activation of the SSMM.  It performed the following main functions:  - initialisation of system controller and control interface    hardware, tables, data, etc.,  - load nominal software from EEPROM to RAM, (reduced) commands    handling, transition to Operational Mode.- The Operational software covering the Operational Mode and Test  Mode. It did run in the MSS RAM and FPC RAM. It performed the  following main functions:  - execution and control of telecommands,  - configuration and test of the memory modules,  - control of data flow from instruments and to TFG to and from the    Memory Modules,  - failure handling, including management of failure log,  - failure recovery,  - creation of event report,  - housekeeping,  - TM packing for all required data, Watchdog control.In case of fatal failure, the SW returns to the Init software toallow for failure investigation.INSTRUMENTS SOFTWARE--------------------Each instrument had its own autonomous SW, located in the instrumentelectronic units. The command and control of the payloads wasperformed by the dedicated Payload Management function of the DMS SW.The physical interface of the DMS PM with the instruments is theRemote Terminal Unit (RTU). Data exchange between the payloads andthe DMS software was performed by means of packetised TM/TC, both forcommands, housekeeping and science telemetry data.-  Commands from the Ground are routed by the DMS software to the   payloads through the RTU and the OBDH bus.-  Housekeeping data from all the instruments are transmitted from   the RTU to the DMS SW through the OBDH bus.-  Scientific data from low rate payloads (PFS, ASPERA, MARSIS,   SPICAM, VMC, OMEGA) are transmitted from the RTU to the DMS SW   through the OBDH bus.-  Scientific data from high rate payloads (OMEGA, MELACOM and HRSC)   are directly transferred to the SSMM through TM packets on the   IEEE-1355 link.GROUND SEGMENT OVERVIEW-----------------------The Mars Express spacecraft will nominally be controlled from the ESANew Norcia (Australia) station during the Routine Operations phase.Shared operations with Rosetta provide a station availability of 8hours a day (design assumption), though longer duration might beachieved during Rosetta cruise phase. Additional Earth stations areconsidered, such as ESA General Purpose Network Kourou 15m stationduring LEOP and NASA DSN 34 m and 70 m stations in critical phases.ESA GROUND SEGMENT- ESA General Purpose Network Kourou station featuring 15 m antennas  with S-band uplink capability and S-band / X-band down-link  capability.- ESA New Norcia station, featuring a 35 m S-band / X-band antenna  with S-band/X-band uplink and down-link capability.DSN COMPATIBILITY- NASA DSN stations featuring 34 m and 70 m antennas, with S-band and  X-band up-link and down-link capabilities, as described in DSN  Flight Project Interface Handbook (NASA/JPL 810.5).Summary of ground stations nominal performances:                                Kourou  New Norcia    DSN       DSN                                15m     35m           34 m      70mS Band Uplink  EIRP (2kW HPA)   81      87            98        117               Pointing Loss    0.05    0.1           0.1       0.1               Antenna Gain     48.5    55.0          55.2      61.7X Band Uplink  EIRP (2kW HPA)   N/A     97            108.8     114.9               Pointing Loss    N/A     0.1           0.3       0.3               Antenna Gain     N/A     64.3          66.8      72.2S Band Downlink G/T at 10 deg   29.85   37.5          40.5      46.9               Pointing Loss    0.03    0.1           0.1       0.1               Antenna Gain     49.2    56.0          56.9      62.3X Band Downlink G/T at 10 deg   38      50.1          50.1      56.7               Pointing Loss    0.1     0.3           0.3       0.3               Antenna Gain     60.0    68.0          68.2      73.1Acronyms--------AOCS        Attitude and Orbit Control SystemHDRM        array hold-down and release mechanismIMU         INERTIAL MEASUREMENT UNITSLV          launch vehicleMLI         multi layer insulationRWA         reaction wheel assemblySADM        solar array drive mechanismSAS         sun acquisition sensorsSUEM        Beagle2 spin-up and ejection mechanismSC          spacecraftSTR         star tracker