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TABLE OF CONTENTS----------------------------------= Instrument Host Overview - Spacecraft= Spacecraft Coordinate System= Mechanical Design - Spacecraft structures - Payload Interfaces= Thermal Control - Thermal Control Concept - Thermal Control design= Mechanisms - Reaction Wheel - Solar Arrays - Beagle-2 - Marsis Antennas Mechanisms= Attitude and Orbit Control System - Star Tracker - Inertial Measurement Units - Sun Acquisition Sensors - Reaction Wheel Assembly - Propulsion Configuration= AOCS Generic Functions - Gyro-Stellar Estimation Function - Reaction Wheel Off-Loading Function - Reaction Wheel Management Function - Thruster Modulation and Selection Function= Propulsion Architecture Description= RF Communications - Overview - Uplink - Downlink= Data Handling Architecture= SSMM Software= Instruments Software= Ground Segment Overview Instrument Host Overview ======================== Data obtained from the Mars Express instruments were send to ground via the spacecraft on-board computer. As spacecraft to Earth communication does typically exclude instrument operations, all data are relayed from the instrument to the spacecraft mass memory, the solid state mass memory (SSMM). The data was downlinked to Earth via the telemetry subsystem using ESA's antenna in New Norcia, Australia, and the Deep Space Network (DSN) antennas of NASA. The radio science experiment required data from the New Norcia and the DSN ground station hardware. This catalogue file gives an overview of the spacecraft and the ground 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 mission customised configuration and mechanical / thermal architecture with a Rosetta inherited avionics. The spacecraft design was driven by mission requirements, science return and system concept. A spacecraft articulation concept with body mounted instruments, fixed High Gain Antenna and 1 degree of freedom steerable Solar Arrays was baselined. The spacecraft design was based on a parallelipedic like shape sizing about 1.7 m length, 1.7 m width and 1.4 m height. The solar array was composed of two wings, providing a symmetrical configuration favourable to aerobraking techniques and minimising torques and forces applied on the arrays and the drive mechanisms during the Mars insertion manoeuvres performed with the main engine. So as to offer the adequate dry mass / propellant mass ratio and large mounting surfaces and volumes for the Orbiter instruments necessary for Mars Express, the traditional cone/cylinder central structure has been found less efficient than a dedicated structural concept with only a Launch Vehicle Adapter connected to a stiffened 'box' as now developed for light weight satellites. Within the overall integrated design of the spacecraft, four main assemblies are planned 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 reaction control system was used for orbit and attitude manoeuvres by either a 400 N main engine or banks of 10N thrusters. The Data Handling is based on packet telemetry and telecommand. The Electrical Power generation was performed by solar arrays, the power storage by a Lithium-Ion battery. A standard 28 V regulated main bus is offered to the payload instruments. The RF Communications function transmitted X Band telemetry 8 hours per day via a High Gain Antenna at rates between about 19 and 230 kbps depending of the Mars to Earth Distance. A variable telecommand rate of 7.81 to 2000 bps was foreseen during up to 8 hour per day. Spacecraft Coordinate System ============================The origin of the spacecraft Reference Frame, named Oa, was located at the separation plane between the spacecraft and the adapter, at the 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 not used at S/C or operations level. Spacecraft Structure and Interface With Payload Units =====================================================The selected SC structure limits the number of complex elements to the bare minimum. Indeed, the only cylindrical part of large dimensions was the Launch Vehicle Adapter ring, the rest of the structural items being principally flat, standard panels with aluminium skins and aluminium honeycomb. The structure was composed of: 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 direction 2) 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 sandwich panels. The panels were made of honeycomb (generally type 1/8-5056- 0.001P) of 10 to 20 mm thickness, bonded to Aluminium facesheets of thickness varying between 0.2 to 0.3 mm and up to 0.5mm additional doubler for local reinforcements. In general, the payload units were accommodated following their main needs. The payloads needing a stringent thermal control and/or pointing performances (HRSC, OMEGA, PFS, SPICAM) were gathered on, or close to, the +X shear wall, inside the spacecraft and close to the AOCS reference (namely the Inertial Measurement Package and the Star Sensors). In order to meet the PFS scanner to PFS sensor co-alignment without disturbance caused by dismounting, these PFS units were installed on a stiff, removable mounting assembly which can be integrated as a single unit on the spacecraft. To expedite installation of the large Omega-SA, this instrument was installed via edge-mounted inserts in the Y-shear wall and a dedicated Omega support panel. The payloads requiring a large field of view and not necessitating stringent thermal control were located externally on the Top and Bottom Floors (ASPERA) or lower edge of the +Y sidewall (MARSIS). None of the payload units required isostatic mountings: they were rigidly fixed to the spacecraft structure utilising standard space- industry inserts and screws. Some of the payload units were of significant mass and therefore require the implementation of large, face-to-face inserts that are bonded inside the sandwich panels at the time of panel moulding. Beagle 2 was accommodated on the top floor of the S/C, in order to minimise dynamic disturbance (centre of mass transfer in the (X, Y) plane) and then maximise the reliability of the Mars orbit insertion manoeuvre. The remaining Beagle 2 hardware after probe ejection is constrained within 50 mm height and is thermally insulated to minimise straylight and thermal distortion disturbances respectively. Thermal Control ===============The spacecraft thermal control was in charge of maintaining all spacecraft equipment within their allowed temperature ranges during all 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 Express spacecraft; it was supplemented with an electrical heating system. The heat rejection toward space was performed using radiators mainly on the +/-Y panels for the platform internal units and the +X panel for the payload equipments. These sides of the spacecraft are the most favourable areas, being most of the time protected from the direct sun inputs (always for the +X side). The Mars planet flux are imposed by the spacecraft orbit and attitude and mainly significant during the pericentre phase in operation. The rest of the spacecraft is insulated with Multi Layer Insulation blankets to minimise the heat exchange and the temperature fluctuations. The spacecraft external units (Platform and Payload units) were thermally decoupled from the spacecraft and provided with their individual radiator when needed. The electrical heater system allowed to raise the temperature of the unit above their minimum allowed limits, with temperature regulation functions provided either by mechanical device or by the onboard software. Most of the spacecraft units were collectively controlled inside defined thermal enclosures in which the heat balances were controlled by proper sizing of heat rejecting radiators and heating power implementation. It allowed to maintain the unit temperatures to acceptable levels. The heat transfer from the units to the radiators was performed by conduction when unit baseplates were attached to the radiator honeycomb panels and by radiation. In that case units and panels had a black finish to maximise heat transfer inside the thermal enclosures. For more demanding units like the HRSC and OMEGA cameras, and the PFS spectrometer, featuring their own thermal control, special precautions were taken by individual trimming of their conductive and radiative isolation. The HRSC camera required a temperature control within a narrow temperature range): it was provided with a thermal strap connecting it to a dedicated radiator tuned to limit the temperature excursion in operation within a 10 degree C temperature range. OMEGA and PFS are provided with dedicated radiators, implemented on the +X side of the Spacecraft. Whatever the Sun / Earth / Mars / Spacecraft geometry, the +X side of the Spacecraft was oriented away from Sun over the complete Martian orbit, both during Nadir pointed science phase and Earth pointed communication phase. This allowed to provide the camera and the spectrometer with a thermal interface at temperature lower than 175K and 190K respectively during the Planet observation. The connection to the radiators were performed by thermal straps, the radiators being themselves decoupled from the rest of the spacecraft using thermal blankets and insulating stand-offs. Payload external units like MARSIS and MELACOM antennas, ASPERA-3 units, were individually controlled units. They required large field of view and thus were directly affected by the external environment and they had to withstand larger temperature ranges than the standard units. They are as far as possible insulated from the spacecraft. Their coatings were selected and trimmed to suit. The spacecraft interface temperature had a very limited influence on their thermal behaviour. The propulsion equipments that were mounted internally were in general isolated with MLI, and provided with their own thermal control heaters: tanks, fluid lines, valves, pressure sensors. The main engine and the thrusters had their thermal coupling with the spacecraft tailored to meet their thermal requirement while preserving the spacecraft thermal behaviour. They were provided with individual electrical heaters sized to maintain these external units within the acceptable temperature range accounting for wide change in radiative environment. The High Gain Antenna was using a passive thermal control: a Kapton/Germanium sunshield was covering the whole antenna on its front side, while a light weight MLI is used on the rear side of the reflector. Mechanisms==========The implementation of mechanisms into the spacecraft configuration had been kept to the minimum. The mechanisms employed are those associated 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 ASSEMBLY The Attitude and Orbit Control System (AOCS) of the spacecraft required implementation of four reaction wheels, used with a three out of four redundancy. They were of ball bearing momentum / reaction wheel type, for clock-wise and counter-clockwise operation, with the wheel mass suspended by two angular contact ball bearings paired by solid preloading. The main functions of the RWA was to ensure correct orientation of the spacecraft in fine pointing modes, and to ensure spacecraft manoeuvrability (e.g. at transition between Mars orbit observation and communication phases), with minimum propellant consumption (the only related consumption lied with wheel momentum off-loading that had to be performed at regular intervals, typically every 2 days.) SOLAR ARRAY DRIVE MECHANISM There were two SADM used on the spacecraft, one for each Solar Array wing. The SADM were mounted on each side of the spacecraft, and were independently controlled by the AOCS Processor Module. The main functions of the SADM was to support the solar array wing throughout the mission, to provide the electrical power and signal interfaces to the spacecraft and to orient the solar array wing towards the Sun by rotation about the Ys axis. The SADM was composed of a motor and gearbox assembly, ensuring the orientation of the solar array by rotation, a shaft and bearing assembly ensuring mechanical connection and pointing accuracy, a twist capsule unit transferring electrical power to the spacecraft. Those elements were mounted on a baseplate which was attached to the spacecraft sidewall. SOLAR ARRAY HOLD-DOWN AND RELEASE MECHANISM Each 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 the structure of each panel which were held together via a stainless steel hold down pin of 3.5 mm diameter on a hold-down baseplate fixed on the spacecraft sidewall. The HRDM also incorporated a pair of pyro initiators, which were actuated after spacecraft separation from the launcher under control of the Data Handling Processor Module. The main functions of the HDRM was therefore to maintain safely stowed each solar array wing and to ensure their release for proper solar array power generation. SOLAR ARRAY DEPLOYMENT SYSTEM Each yoke and wing of the solar array was fitted with a deployment mechanism that ensured proper deployment and latching of the solar array after release of the HDRM. The deployment mechanism consisted in a set of spring energy driven hinges mounted by pair between each solar array panel, between the first panel and the yoke, and between the yoke and the SADM. Each hingeline was then linked to the others by a set of pulley and cables, that ensured a synchronised deployment of the wing. The torque margin of the Solar Array deployment system varied between 7.5 (at beginning of deployment) and 2.6 (at end of deployment). BEAGLE 2 SPIN-UP AND EJECTION MECHANISM Beagle 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 the probe, which interfaced to the orbiter top floor through the Spin-Up and Ejector Mechanism. MARSIS ANTENNAS DEPLOYMENT MECHANISMS The baseline configuration for MARSIS deployment mechanism had departed from the Cassini (STEM) concept, i.e. a tubular antenna made of 2 semi-circular formed strips made of Copper-Beryllium. The selected design was the ASTRO one, consisting of a boom made of a GFRP tube pierced with 2 diametrically opposed diamond shaped holes at the selected distance to provide folding capability. The antenna boom contained two wire elements forming the active radioelectrical part of the antenna, and was folded at each hollowed hinge and held flattened in specific containers. When release was initiated, the container was opened through pyro devices, and the boom was self deploying thanks to its intrinsic energy which had been stored during the folding/flattening process necessary to meet the launch volume constraints. Attitude and Orbit Control System================================= AOCS BASIC CONCEPTS Due to the selection of a fixed high Gain antenna (HGA), and to the propulsion configuration including a Main Engine, the Mars Express mission required a high level of attitude manoeuvrability for the spacecraft. 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 on ground, using a polynomial description of the Quaternion to be followed by the Spacecraft. The attitude estimation was based on Star Tracker and gyros, ensuring the availability of the measurements in almost any attitudes. Some constraints had however to be fulfilled, the Star Tracker being unable to provide attitude data, when the sun or the planet are close to or inside its Field of view. Reaction wheels were used for almost all the attitude manoeuvres, providing a great flexibility to the Spacecraft and reducing the fuel consumption. The angular momentum of the wheels had however to be managed carefully from ground. STAR TRACKER (STR) The Star Tracker (STR) was the main optical sensor of the AOCS, used at the end of the attitude acquisition to acquire the final 3-axes pointing, and during almost all the nominal operations of the mission. A medium Field Of View (16.4 deg circular) and a sensitivity to Magnitude 5.5 were used to provide a 3-axes attitude measurement with at least 3 stars permanently present in the FOV. The STR included a star pattern recognition function and can perform autonomously the attitude acquisition. The Mars Express Star Tracker was produced by Officine Galileo, and is similar to the Rosetta one, except at S/W level. 2 Star Trackers were implemented on the minus Xa face of the Spacecraft, with an angle of 30 degree between their optical axes. INERTIAL MEASUREMENT UNITS (IMU) Two Inertial Measurement Units (IMU) were used by the AOCS, each IMU including a set of 3 gyros and 3 accelerometers aligned along 3 orthogonal axes. The AOCS control used either the 3 gyros of the same IMU (reference solution at the beginning of life) or any combination of 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 to the lower criticality of the accelerometer function, and to the availability onboard of an alternative method for the delta V measurement (pulse counting). The Gyros were useful during the attitude acquisition phase for the rate control, during the observation phase to ensure the required pointing performances and during the trajectory corrections, for the control robustness and failure detection. A non mechanical technology was selected to avoid the mechanical sources of failure in flight. The Accelerometers were essential during the main trajectory corrections such as the insertion manoeuvre to improve the accuracy of the delta V. The IMU of Mars Express is identical to the Rosetta unit. Only the number of units and the onboard management of the configuration was different. SUN ACQUISITION SENSORS (SAS) Two redunded Sun Acquisition Sensors (SAS) were implemented on the Spacecraft central body and are used for the pointing of the Sun Acquisition Mode (SAM) during the attitude acquisition or reacquisition in case of failure. The SAS are identical to Rosetta units, 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 to perform most of the nominal operations of the mission with a 3 RWL configuration among 4. During some critical phase during which the transition to the SAM had to be avoided (before lander ejection and before Mars Insertion Manoeuvre), a 4 wheels configuration was be used, under ground request. The Reaction wheels provided the AOCS control torques during all the phases of the mission except the trajectory corrections, the attitude acquisition and back up modes. PROPULSION CONFIGURATION The Propulsion configuration included a Main Engine (414 N) which was used to perform all the major trajectory changes, and 10 N thrusters used for the attitude control and also to produce the thrust during the small trajectory corrections. The 10 N thrusters configuration was optimised to perform all the attitude control functions with only 4 redunded thrusters, each of them being implemented near a corner of the -Z face of the spacecraft. SOLAR ARRAY DRIVE MECHANISM 2 redunded Solar Array Drive Mechanisms (SADM) were implemented on the Y+ and Y-walls of the spacecraft to control the orientation of the Solar Arrays. The SADM was only used for large angle orientation of the wings, the selected flight orientation during the observation phase near pericentre requiring no SADM actuation, once the observation attitude was reached. The SADM used a stepper motor, a gear, and a twist capsule technology. The SADM motion is defined in the range +/-180 deg (minus margins). The SADM is identical to the Rosetta unit, except for the speed levels which are specific to Mars Express. AOCS HARDWARE ARCHITECTURE AOCS unit Nb Technology / characteristics Heritage Supplier ------------ ---------------------------- ---------- -------- Star Tracker 2 CCD detector. 16.4deg Rosetta unit. Officine circular FOV/ Magnitude 5.5 Galileo Gyro/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-TNO Sensor (SAS) a pyramid Reaction Wheel 4 Ball bearing Momentum/ Telecom. Sat. Teldix Reaction wheels. Unit 12 Nms/0.075 Nm SADM 2 Stepper motor with gear. Rosetta unit Kongsberg Twist capsule AOCS Generic Functions======================The AOCS modes used generic functions for the guidance, the attitude estimation and the actuators management. The role of the guidance was to provide onboard the reference attitude to be followed at each time of the mission by the attitude control. It concerned of course the orientation of the Spacecraft but also the Solar Array position. The analysis of the mission needs showed that 4 types of guidance are necessary 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 ephemerides (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 software level : - 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 FUNCTION The gyro-stellar estimation function was common to many AOCS modes : It was initialised during the Sun Acquisition Mode (SAM) to prepare the following Earth acquisition operation (SHM: Safe/Hold Mode). It provided accurate attitude estimation during the Normal Mode of course but also in the Orbit Control Mode (OCM) and Thruster Transition Mode( TTM) for instance. The gyro-stellar estimator processed gyro and star tracker (STR) measurements to provide an accurate estimate of the spacecraft attitude. It was based on a Kalman filter with constant covariance that allowed mixing measurements at different rates (8 Hz for the gyros and 2 Hz for the STR). The constant covariance reduces the computer load while ensuring good performances. The estimated attitude was a quaternion representing the spacecraft attitude in the J2000 inertial frame. The gyro-stellar estimator also estimated the gyros drifts to limit the attitude errors in case of STR measurement absence due, for instance, to a temporarily STR occultation. A specific management of the drift estimates was proposed for Mars Express, taking into account the specific conditions of the scientific mission phase (existence of rates due to varying profiles, and potential occultation). The gyro-stellar estimator implemented a coherency test between the gyro and STR measurements in order to detect failures that could not be detected at equipment level. REACTION-WHEEL OFF-LOADING FUNCTION The wheel Off-Loading function enabled to manage the angular momentum of the wheels to a target value, through thruster actuations. This function was completely autonomous during the last phase of the Earth acquisition sequence (SHM/EPP:Earth Pointing Phase). During the nominal operations around Mars, it was preferable to command the wheel Off-Loading from the ground, the date being optimised taking into account the mission constraints. The Off-Loading function managed simultaneously all the wheels. It included several sequences of thruster pulses until angular momentum of each wheel was close to the target value. This sequence was defined by a feed forward 3-axes wheel torque command combined with a thruster pulse. The sequence ended with a tranquillisation phase controlled by the wheels, in order to damp the dynamic excitation generated by the actuation of thrusters and wheels. REACTION WHEEL MANAGEMENT FUNCTION This function was active in all the modes controlled through wheel torques (Normal Mode and Safe/Hold Mode at the end of the attitude acquisition sequence), but also when the wheels were kept to a constant speed through a specific control loop but not used in the AOCS control, as in Orbit Control Mode, Thruster Transition Mode or Braking Mode. Six states of the wheel configuration are possible with this 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 fourth wheel if a hot redundancy is useful (4t). During trajectory corrections the configuration included 3 wheels controlled in speed (3s). Intermediate states are necessary between these basic configurations in order to spin the wheels for instance (3t + 1s). This function was also in charge of the generation of wheel torque commands in wheel frame, and of the friction torque estimation necessary for compensation and for the failure detection. It interfaced also with the Wheel Off-Loading function. THRUSTER MODULATOR AND SELECTION FUNCTION The selected amplitude modulator and on-time summation algorithms were re- used from Rosetta and adapted to match more efficiently the Mars Express needs taking into account the specific thrusters configuration. The modulator had only one working phase where the four thrusters can be 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 modulation type used (ON-modulation or OFF-modulation) was automatically selected so as to maximise the available torque capacity for attitude control. In the case the torque capacity was insufficient with respect to the commanded control torque, priority is given to the control and the commanded force ratio is automatically modified to recover the required torque capacity. Moreover in order to limit the actuation delay, the attitude control torque was always produced at the beginning of the actuation period. To limit the number of thrusters ON/OFF or to tune the control limit cycle amplitude when using thrusters, the modulator output period had to be changed to any period multiple of 125 ms. Propulsion Architecture Description===================================A bi-propellant system based on a telecommunication spacecraft heritage was adopted for the baseline. A set of isolation pyro valves and latch valves had been added to ensure safe operations during Launch and Cruise, and for a re- liable acquisition of the Mars orbit for science mission. At launch, the pressurant assembly (high and low pressure sections) were all isolated from the propellant tanks by normally closed pyrotechnic valves PVNC1 to PVNC6, by non return valves NRV1 to NRV4. The propellant tanks are pressurised to 4 bar. Similarly, the propellant was isolated from the Reaction Control Thrusters and Main Engine assembly by normally closed pyrotechnic valves PVNC7 to PVNC14 and thruster/main engine Flow Control Valves (FCV). Following separation, the pyro valves protecting the pressurant assembly were fired to pressurise the system to its operating pressure of 17 bar. Then the latch valves were closed, isolating the non return valves from propellant. A pressure transducer (PT2) located at the regulator outlet could monitor pressure build up at the NRV location due to regulator leakage. When necessary the latch valves were opened and the pressure relieved into the propellant tanks. It was assumed that a pressure of up to 20.5 bar could be the criterion to initiate an open/ close cycle of the latch valves by telecommand. The 20.5 bar pressure was an initial suggestion which needed to be confirmed. It may affect component qualification issues because it exceeds existing MEOP values for the components in the section. Short duration opening times for the latch valves minimised propellant vapour migration and it was essential for both oxidiser side and fuel side latch valves to open simultaneously to limit vapour migration. The system operates in this pressure regulated mode, using the 10 N Reaction Control Thrusters only, during the period of the transfer orbit to Mars. A few days before Mars orbit insertion, the 400 N Main Engine was primed and then calibrated by specific blank manoeuvres, combined with re-targeting of the S/C after Beagle 2 probe ejection. This ensured that the Main Engine could be used safely for the Mars orbit insertion and acquisition of the operational orbit. Should a Main Engine failure be detected at this stage, a back-up scheme, using the Reaction Control Thrusters would have been implemented to reach at least a degraded orbit around Mars. After attaining the operational orbit, the pressurant and Main Engine assemblies were re-isolated by firing all the normally open pyrotechnic valves and closing the latch valves. The remainder of the mission was per- formed in blow down mode, using only the 10 N Reaction Control Thrusters. The number of Reaction Control Thruster had been limited to 8 (4 nominal, 4 redundant), located at the bottom (-Z) side of the spacecraft to provide thrust principally along Zb to compensate for Main Engine thrust imbalance caused by Main Engine alignment and Spacecraft Centre Of Mass (CoM) uncertainties. Adequate tilting of the Reaction Control Thrusters is implemented so as to provide the capability for torque around each main axis of the spacecraft. 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, permitting individual switch over from prime to redundant for each Reaction Control Thruster. It had to be noted that this two-tank configuration was compatible with a horizontal handling of the spacecraft as required by Soyuz launch campaign, on the proviso that the tanks were filled at least up to 62% of their maximum capacity. The compatibility of this fill fraction wrt S/C global dynamic behaviour was under investigation to avoid fluid/structural modes coupling. RF Communications================= OVERVIEW The communications with the Earth could be performed either in S-Band or X-Band in accordance with ESA Standards. Two Low Gain Antennas (LGA) allow omni-directional emission and reception in S-Band, while a dual band 1.65 m High Gain Antenna (HGA) allows high rate TM emission 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 Dual Band Transponder before routing the resulting bit flow to the Data Handling. The stored TM within the SSMM is modulated in either SBand or X-Band within the Dual Band Transponder, which also performed S- Band signal amplification with 5 W. X-Band signal amplification is performed using a 65 W Travelling Wave Tube Amplifier. UPLINK The communication from the ground station(s) to the spacecraft was performed 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 Gain Antenna and the other one on the bottom of the spacecraft, thus allowing a quasi omnidirectional coverage. The LGA was used mainly during Launch and Early Operation Phase (LEOP), critical phases and for emergency situations. A narrow-beam dual-band high-gain antenna was 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 packetised telecommands as NRZ/PSK/PM data, was routed towards a diplexer, performing frequency discrimination, and then to the Dual Band Transponder input. The transponder performed carrier acquisition and demodulation, and transmitted the extracted signal to the Data Handling 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 Express Spacecraft 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 are selectable by Memory Load Command (MLC). As a baseline, the lowest bit rates was used in case of emergency via the Low Gain Antennas in S-Band, while the highest ones were used operationally through the High Gain Antenna in XBand. DOWNLINK A high data downlink capability was required, considering the large data volume generated by the instruments. Nevertheless, downlink capacity was limited by the large spacecraft to Earth distance. The downlink of the telemetry data to the ground stations were performed in 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 CDMU design allowed to generate a telemetry flow at any bit rate corresponding to a power of two multiplied by 32/n and lower than 262.144 bps, where n is equal to 2, 3, 5 or 7. The possible bit rates were selected via Memory Load Command (MLC) and vary from 7.8 bps as a minimum and can be up to 230 kbps. The bit rate to which reference was made was the bit rate following Reed-Solomon encoding, but prior to convolutional encoding, if any. Due to hardware limitations, convolutional encoding was only performed for bit rates lower than 65536 bps. Above this value, only Reed-Solomon encoding was performed. As a baseline, the lowest bit rates were used in case of emergency only using the Low Gain Antennas, whilst the highest ones were used operationally through the High Gain Antenna in X Band. The variable bit 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 S Band or X Band by the Dual Band Transponder, and added to the MPTS ranging signal if it had been detected on the uplink. Data Handling Architecture==========================The Data Management System (DMS) was in charge of telecommand distribution to the whole spacecraft, of telemetry data collection from the spacecraft sub- systems and payloads and data formatting, and of the overall supervision of spacecraft and payload functions and health. The DMS was based on a standard OBDH bus architecture enhanced by high rate IEEE 1355 serial data link between the CDMU (Control and Data Management Units) processors and the SSMM and STR. The OBDH bus was the data route for platform and payloads data acquisition and commands distribution via the RTU. The DMS included 4 identical Processor Modules (PM, 1 to 4) located in the 2 CDMU. Two processor modules were dedicated to the DMS (PM2 and PM3), and two to the AOCS(PM1 and PM4). The PM selected for the DMS function acted as the bus master. It was in charge of Platform subsystem management (Communications, Power, Thermal, Payloads). The PM selected as the AOCS computer was in charge of all sensors, actuators and Solar Array Drive Electronics (SADE). TC-decoder and Transfer Frame Generator (TFG) were included in each CDMU. The Solid State Mass Memory (SSMM) was used for data storage including 12 Gbits of memory at BOL. It was coupled to the two DMS processors, the TFG, OMEGA, HRSC and MELACOM instruments. It stores science and global housekeeping telemetry packets. OVERVIEW The Data Handling architecture was organised around the two CDMU. They were in charge of controlling ground command reception and execution, on-board housekeeping and science data telemetry storage and formatting them for transmission. The on-board data management, controlled processing and execution of on-board control procedures belongs to their tasks as well. Each CDMU featured two MA3-1750 Processor Modules, each of them being able to process either Data Management or AOCS software. A built-in failure operational Reconfiguration Module (RM) ensured system level FDIR and reconfigured the CDMU as necessary. Data transfer with other Data Handling units were ensured using standard links such as a redunded OBDH data bus or IEEE-1355 serial links. Two Interface Units were performing inter- face adaptation between those links and other spacecraft units. The AOCS Interface Unit (AIU) was dedicated to AOCS equipment, while the RTU interfaces with the remainders, including the Instruments. A file-organised 12 Gbits SSMM was implemented to store the Housekeeping and the Science Data. It also collected directly Science Data from the three high rate Payload Instruments. 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 the micro- controller located in the FPC. The main part of the SSMM-SW is programmed in C language. Parts of the start-up function are programmed 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 to allow for failure investigation. Instruments Software====================Each instrument had its own autonomous SW, located in the instrument electronic units. The command and control of the payloads was performed by the dedicated Payload Management function of the DMS SW. The physical interface of the DMS PM with the instruments is the Remote Terminal Unit (RTU). Data exchange between the payloads and the DMS software was performed by means of packetised TM/TC, both for commands, 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 ESA New Norcia (Australia) station during the Routine Operations phase. Shared operations with Rosetta provide a station availability of 8 hours a day (design assumption), though longer duration might be achieved during Rosetta cruise phase. Additional Earth stations are considered, such as ESA General Purpose Network Kourou 15m station during 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 70m S 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 61.7 X 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 72.2 S 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 62.3 X 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 73.1 Acronyms -------- AOCS Attitude and Orbit Control System HDRM array hold-down and release mechanism IMU INERTIAL MEASUREMENT UNITS LV launch vehicle MLI multi layer insulation RWA reaction wheel assembly SADM solar array drive mechanism SAS sun acquisition sensors SUEM Beagle2 spin-up and ejection mechanism SC spacecraft STR star tracker
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