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6. SPACECRAFT COMPOSITE - ESA

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6. SPACECRAFT COMPOSITE6.1 SYSTEM CONFIGURATIONSThe satellite design concept is thoroughly modular, in order to accommodate without conflict therequirements arising from the various environments and mission scenarios. The design comprises twopropulsion modules and three scientific modules:• Solar Electric Propulsion Module (SEPM) for the interplanetary transfer• Chemical Propulsion Module (CPM) for Mercury capture and orbit acquisition• Magnetospheric Orbiter (MMO) for the particle-and-field experiments• Planetary Orbiter (MPO) for the remote sensing and radio science experiments• Surface Element (MSE), a lander for short-lived in-situ measurements.The five modules are combined in different ways according to the requirements of each mission scenario.The scientific modules (MMO, MPO, MSE) are designed to fit either the single-launch or the dual-launchscenario with minimal changes. The propulsion modules (SEPM, CPM) for the single-launch scenario areroughly twice the size as those designed for the dual launch option. In the dual launch option, the propulsionmodules for the first and second launch are made identical, by taking the worst-case requirements into account,thus offering the advantages of a recurrent module procurement.Single launch scenarioIn the single-launch scenario, all modules are delivered by one dedicated launch of Ariane-5. The modulesare stacked to be deployed in the following order, and with the following separation interfaces:• jettison the Electric Propulsion Module before capture (separation interface between SEPM and CPM)• deploy the Magnetospheric Orbiter, after MMO orbit insertion manoeuvre (spin-eject device, mounted toCPM)• jettison CPM, after MPO orbit insertion manoeuvre (separation interface between CPM and PlanetaryOrbiter)• deploy the Surface Element, after acquisition of MPO orbit (spin-eject device, mounted to MPO).The propulsion elements include four (plus one cold-redundant) 200 mN thrusters (SEPM) and one 400 Nbipropellant thruster plus eight 20 N thrusters for attitude control (CPM). SEPM can accommodate either ion orSPT thrusters; the chief differences are in the required propellant mass and the solar array size (6.5 kW for theSPT thrusters and 10.5 kW for the ion thrusters). The solar array employs GaAs cells with 20% optical solarreflectors, in two wings. Through the interplanetary and early Mercury orbit phase, the command and controltasks are centralised in the Planetary Orbiter, providing overall monitoring and control tasks, and managingtelecommunications to Earth. In Mercury orbit, MPO and MMO are independent, and the MSE data relay ismanaged by MPO.Split launch scenarioIn the dual-launch scenario, a first launch of a Soyuz-Fregat vehicle delivers a composite made ofSEPM+CPM+MMO+MSE, while another launch delivers a composite consisting of SEPM+CPM+MPO.The propulsion elements feature three 200 mN ion thrusters (SEPM) and one 400 N bipropellant thruster pluseight 20 N thrusters for attitude control (CPM). The SEPM features a 5.5 kW GaAs cell solar array (with 20%OSR), in two wings. Through the interplanetary and early Mercury orbit phase, the command and control tasksare centralised in the respective Orbiters. MMO manages the MSE data relay in the first of the two dual-launchmissions.The system configurations in the two mission scenario options are described in more detail below.6.1.1 Single Launch ConfigurationThe spacecraft composite configuration in the single-launch design is driven by the accommodation of thefive modules on Ariane-5 with a short fairing (including the structural requirements), by the thermal andoperational conditions during the cruise, and by the requirements for staged release of the modules at theappropriate times during the mission.113


BepiColomboThe composite (Figure 6.1-1) interfaces with the launcher via a standard 1666 mm diameter interface ring.The main structure is constituted by the SEPM thrust cone; the CPM cone structurally continues the SEPM conevia another separation interface. An ad-hoc structure further connects the main structure to the Planetary Orbiterstructure, and supports the third separation interface. The height of the composite is made as short as possible bymaking the nozzle of the 400 N CPM engine protrude into the SEPM down to the level of the Xenon tanks(Figure 6.1-2). SEPM is a box structure containing the Xenon tanks, inside the thrust cone, and the power andregulation equipment. The folded solar panels are placed symmetrically on two sides of the box. In order to limitthe box height, the upper fixation points of the solar array are provided by a rod structure. The size of the SEPMhas been designed to fit both SEP thruster options, with their different propellant loads.CPM is a simple structure containing two bipropellant tanks and the pressurant tank. The MagnetosphericOrbiter is mounted on one side of CPM (the cold side during the cruise); the interface is a spring-eject device.MSE is structurally attached to the Planetary Orbiter via another spring-eject device, and is thermally protectedin a dedicated enclosure during the cruise.The mounting of the Orbiter to the composite is dictated by thermal and propulsion reasons. Once separatedfrom SEPM, the remaining composite (Planetary Orbiter + CPM with Magnetospheric Orbiter and MSEattached) must survive in the close proximity of the planet for the duration of the CPM manoeuvres. Hence, theattitude of the Orbiter must be the nominal one, with the +Z side towards the planet and the X-axis parallel tothe direction of motion. Therefore, the CPM nozzle must be in the X direction of the Orbiter, that is, mounted onone side (Figure 6.3-2). This ‘vertical’ mounting of the Orbiter in the composite complicates the structuraldesign in the interface area, and will need optimisation in future work. The orientation of the Orbiter in thecomposite is with the ‘hot’ (antenna) side towards the Sun during the cruise. MSE is mounted on the +Z (planet)side of the Orbiter (the ‘cold’ side during the cruise), with its own device on a protruding platform for anejection along the X direction.Table 6.1-1 presents the mass budget summary of the launch configurations in 2009 with the ion thrustersand the SPT thrusters, the most demanding option (Table 5.1-5). The availability of an Ariane 5 with a restartableupper stage (Ariane 5V) has been assumed, which provides an adequate mass margin (>20 %).Summary Launch Mass Budget [kg]2009 launchIon Thrusterswith GridsMagnetospheric Orbiter 157.2 157.2Surface Element 234.3 234.3Planetary Orbiter 357.3 357.3Chemical Propulsion Module dry 167.4 167.4Subtotal 1 (dry mass at Mercury) 916.1 916.1Bipropellant 335.9 335.9Subtotal 2 (mass after jettison) 1252.0 1252.0SEPM Dry 664.1 621.5Subtotal 3 (mass before jettison) 1916.1 1873.5Cruise propellant 474.8 803.2Launch vehicle adapter 92.0 92.0Launch mass 2482.5 2768.7Ariane 5 limit launch mass 3500 3500System margin (kg) 1017.1 731.3System margin (%) 41.0 26.4Table 6.1-1: Single launch mass budget summary (2009)SPTThrusters114


Spacecraft CompositeFigure 6.1-1: Single-launch configuration underthe Ariane 5 fairing.Figure 6.1-2: Cross section of the launchconfiguration.Figure 6.1-3: 3D view of the launchconfiguration.Figure 6.1-4: Cruise configuration.115


BepiColombo6.1.2 Split Launch ConfigurationThe configuration for the split launch option must fit the fairing of the Soyuz-Fregat vehicle, i.e.:• Launch 1: SEPM+CPM+MMO+MSE• Launch 2: SEPM+CPM+MPO.The composite interfaces with the launcher via a 945 mm diameter (Mars Express) interface ring. Thepropulsion elements are designed to be nearly the same in the two launches. SEPM is built up according to thesame criteria as in the Ariane launch, with reduced size due to the smaller propellant load and solar array. TheSEPM Xenon tanks and solar array are sized according to the ion-thruster-with-grids requirements, as the SPTthrusters do not provide enough mass margins to fit the Soyuz capability. The CPM for launch 1 includes aspecial interface structure to accommodate the MSE-MMO stack, and additional equipment (star sensors, solararray) that are not needed in launch 2. The propellant loads and tanks are adapted to the different requirements(cruise and insertion ∆v) of the various options. MMO and MPO are identical to the single-launch configuration.In MSE, the insertion motor is removed, as the insertion manoeuvre is performed by CPM.Figures 6.1-5 and 6.1-7 show the launch and flight configuration for launch 1, and Figures 6.1-6 and 6.1-8 thecorresponding configurations for launch 2. Table 6.1-2 shows the mass budgets for launches from Baikonur(2007/2008 and 2009). Discrepancies in launch masses are partly due to small differences in cruise and insertion)v. Kourou (latitude 5°N) is better suited for the 2009 launches with a low asymptote declination (5°S). Shouldthe pad be available in 2009, the available launch mass would increase significantly, as shown in Table 6.1-2.CompositeLaunch yearMMO+MSE2007/2008MMO+MSE2009MPO2009Magnetospheric Orbiter 157.2 157.2 0.0Surface Element 219.5 219.5 0.0Planetary Orbiter 0.0 0.0 357.3Chemical Propulsion Module dry 89.4 89.4 77.1Subtotal 1 (dry mass at Mercury) 466.1 466.1 434.4Bipropellant 119.9 125.8 158.3Subtotal 2 (mass after jettison) 586.0 591.9 592.6SEPM Dry 365.5 365.5 365.5Subtotal 3 (mass before jettison) 951.4 957.4 958.1Cruise Propellant 255.7 257.8 255.8Launch vehicle adapter 49.0 49.0 49.0Launch mass 1254.1 1265.3 1265.7Baikonur launch siteSoyuz-Fregat limit launch mass 1330.0 1330.0 1330.0System margin (kg) 75.9 64.7 64.3System margin (%) 6.0 5.1 5.1Baikonur launch site + lunar swing-bySoyuz-Fregat limit launch mass 1550.0 1550.0 1550.0System margin (kg) 295.9 284.7 284.3System margin (%) 24.0 22.5 22.5Kourou launch siteSoyuz-Fregat limit launch mass 1510.0 1510.0System margin (kg) 244.7 244.3System margin (%) 19.3 19.3Table 6.1-2: Split-launch mass budget summary (kg)116


Spacecraft CompositeFigure 6.1-5: Split-launch 1 (MMO + MSE);configuration under the Soyuz fairing.Figure 6.1-6: Split-launch 2 (MPO);configuration under the Soyuz fairing.Figure 6.1-7: Split-launch 1 (MMO+MSE);flight configuration.Figure 6.1-8: (left) Split-launch 2 (MPO); flightconfiguration.Figure 6.1-9: (right) Split spacecraftconfiguration under Ariane 5 fairing.It has been verified that the split-spacecraft configuration is also compatible with a single Ariane 5 launch,using the Speltra adaptor (Figure 6.1-9). The launch window is the same as for the baseline Ariane 5 scenariodescribed in Section 6.1.1; the mass margin remains larger than 20 %.117


BepiColombo6.1.3 Concluding Remarks on System ConfigurationThe current study results demonstrate the ample flexibility of launch options compatible with the missionrequirements, and the capability of the mission to be adapted to varying programmatic constraints.Results have been obtained for three alternative system configuration options (single-launch, and split-launchwith or without lunar swing-by), the advantages and drawbacks of which are summarised in Table 6.1-3.SystemConfigurationOptionsAdvantagesDrawbacksSingle-LaunchScenario(Ariane 5)Split-LaunchScenario (Soyuz-Fregat)Split-LaunchScenario withLunar Swing-by(Soyuz-Fregat)• Comfortable system mass margins(>20%) for 2009, 2010 launchopportunities• Possibility of using either ion or SPTthrusters, which is advantageous withrespect to the commercial availabilityof either type of thruster• Possibility of launching 2 separatespacecraft (as designed for split-launchscenario) on a single Ariane 5• Lower cost of Soyuz-Fregat launches• Early launch (of MMO, MSE) possiblein May 2007, arrival at Mercury inAugust 2009• Comfortable system mass margins(>20%) for 2008, 2009 launchopportunities from Baikonur• Low cost of Soyuz-Fregat launches• Higher cost of Ariane 5 launch• Very low system mass margin (5%)for 2007, 2009 launch opportunitiesfrom Baikonur• Early launch in 2007 requires fasttechnology development actions to beinitiated• Less flexibility in the choice of theelectrical thrusters• Cruise time to Mercury increases from2.5 to 3.5 years• Early launch in 2008 requires fasttechnology development actions to beinitiated• Less flexibility in the choice of theelectrical thrustersTable 6.1-3: System configuration options advantages and drawbacksIn the case of the proposed lunar swing-by strategy (Section 5.1.4), the launch capability of Soyuz-Fregatincreases from 1,330 kg to 1,550 kg, providing the required >20% margin, at the cost of a one-year longertransit time to Mercury. In addition, several other possibilities for increasing the mass margin in a split launchinclude the following:• Given the assumption of a common procurement for both launches, the CPM is not optimised for either. Adedicated design could reduce the system launch mass.• The specific impulse of ion thrusters used for this study has been the end-of-life value, corresponding to10000 h life time. Using a more realistic value (corresponding to the thruster performance at the actual endof mission, i.e. in the order of 7000 h) will reduce the required propellant mass.• The CPM could be merged with the MSE propulsion system in Launch 1. This approach is pursued in thealternative MSE design presented in Section 6.6.2.• An in-depth mass reduction exercise could be carried out across all mission elements.• The Soyuz-Fregat performance in 2007, and even more so in 2009, may be improved with respect to thepresent manifest. In particular, if a launch facility in Kourou is developed for this launcher, the launchcapability for an escape at 2.3 km/s increases to 1,510 kg, which would also provide the required 20%margin (Table 6.1-2). However, a launch from Kourou cannot be taken as the mission baseline at this time,since the availability of a Soyuz launch pad in Kourou depends on a commercial decision, planned but notyet finalised.118


Spacecraft CompositeIn case the mass margin of at least one of the two composites can be recovered by some of the aboveapproaches, a split-launch scenario can be envisaged in which the heavier composite is launched in the lunarswing-by window, and the lighter composite one year later without the lunar swing-by, in which case they willboth reach Mercury approximately at the same time (this could be of interest for science operations). However,this would require two different sets of mission operations to be developed (with and without lunar swing-by).Table 6.1-4 shows a summary of possible combinations of split-launch scenarios, compared with two singlelaunchscenarios. Those which are assumed as baseline scenarios at the end of this study are shown in boldcharacters.Payload Launch Launch Date Arrival at MercuryHeavy Composite(MMO + MSE)Soyuz-Fregat, Baikonur+ lunar swing-byJanuary 2008August 2011Heavy Composite(MPO)Soyuz-Fregat, Baikonur+ lunar swing-byAugust 2009October 2012Heavy CompositeSoyuz-Fregat, Baikonur+ lunar swing-byJanuary 2008August 2011Light CompositeSoyuz-Fregat, BaikonurJanuary 2009August 2011Heavy CompositeSoyuz-Fregat, Baikonur+ lunar swing-byAugust 2009October 2012Light CompositeSoyuz-Fregat, BaikonurAugust 2010October 2012Single Composite Ariane 5, Kourou January 2009 August 2011Single Composite Ariane 5, Kourou August 2010 October 2012Table 6.1-4: Possible launch combinations (baseline scenarios in bold characters)6.2 SOLAR ELECTRIC PROPULSION MODULE6.2.1 Module ConfigurationThe Solar Electric Propulsion Module (SEPM) configuration depends on the thruster type adopted (the ionthrusters require a smaller propellant load and a larger solar array than the SPT thrusters; see Section 6.2.2below) and the launch scenario.Ion ThrustersSingle launchSPTThrustersSplit launchIon ThrustersPropulsion subsystem 204.2 219.2 117.7Solar Array and SADM 224.2 166.6 128.7Power and harness 36.0 36.0 18.0Thermal control 34.6 34.6 17.3Structure and mechanisms, incl. interfacestructures165.1 165.1 83.8SEPM Dry 664.1 621.5 365.5Cruise Propellant 474.8 803.2 255.7Total wet mass 1138.9 1424.7 619.2Table 6.2-1: SEPM mass budget summaries. Mass items include margins according to item maturityThe general features of SEPM are as follows. The module is a simple rectangular prism. A central thrust coneis the main structural element, and transmits the loads to the launcher interface. The Xenon propellant tanks arehoused within it. The tank volumes have been selected, to suit the configuration constraints, from existing highpressure tanks (assumed Xenon storage pressure 125 bar @ 50 C); Xenon tanks of large volume are not119


BepiColomboavailable today. The thruster number and configuration, and the solar array size, depend on the launch scenario.In the single launch option, four 200 mN thrusters are installed at the level of the launcher separation plane, atthe corners of a square. An additional thruster is installed at the centre, to produce a fully one-failure-tolerantconfiguration. The thrusters are clustered near the centre to minimise misalignment torques. Two solar arraywings, with multiple panels, provide the cruise power (ion thruster option: 60 m², 10 kW at 1 AU, five panelsper wing; SPT thruster option: 48 m², 8 kW at 1 AU, four panels per wing). On each wing, a solar array drivemechanism (SADM) provides one rotational degree of freedom, around the yoke axis, sufficient for the arrayorientation needs during the cruise. The propulsion drive and power conditioning electronics equipment aremounted on the bottom platform. There are three thrusters in the dual launch configuration; ion thrusters areemployed to provide sufficient mass margins. The solar array wings have five panels each, smaller than in theAriane options (33 m², 5.5 kW at 1 AU).Table 6.2-1 shows the SEPM mass budget summary. Figure 6.2-1 shows a 3-D view of the module.(kg)6.2.2 Electric propulsion120Figure 6.2-1: SEPM 3-D view.Electric propulsion systems generate the thrust by acceleration of a propellant by electric energy. The exhaustvelocity is in general higher than for chemical propulsion systems and the propellant necessary for a dedicatedmission decreases accordingly. A break-even point in favour of electric propulsion is normally achieved atsatellite masses above 1000 kg and ∆v > 500 m/s, which is certainly the case for an interplanetary mission. Thesaving of propellant mass is partly counteracted by a relatively high dry mass of the electric propulsion system(thrusters and power conditioning units), and the power source. For the Mercury mission, requiring ∆v > 6km/s, the following criteria for the selection of a thruster system were applied: thrust level above 100 mN;specific impulse above 1500 s; European supplier preferred. Three thruster types meet the above mentionedcriteria.SPT-type thruster, 140 mm discharge chamber diameterSPT (Stationary Plasma Thrusters) emerged from a Russian propulsion technology a few years ago, with theheritage of frequent flight experience on Russian spacecraft. At present the SPT 100 (100 mm ioniser diameter,80 mN thrust, 1500 s Isp) is under consideration for application to North/South station keeping of manygeostationary satellites. A formal qualification to western standards and performances, required by telecomspacecraft, has been done by ISTI, a joint venture between SS/L, FAKEL, ARC, RIAME and SEP. The PPS1350 developed by SEP (maximum thrust level 83 mN) is foreseen for the application to SMART-1. In additionwork is going on in USA, Russia and Europe (MMS and SEP) to develop the higher thrust SPT 140 engine (140mm ioniser diameter) for orbit transfer to GEO and for interplanetary missions.Electron bombardment Ion Thruster, 220 mm beam diameterElectron bombardment ion thrusters are manufactured by MMS and DERA, England (T6 IPS). Developmentactivities over many years in the UK have culminated in the development of a 10 cm beam diameter gridded ionthruster (UK 10, also referred to as T5) for tasks requiring a moderate thrust below about 30 mN. The UK 10will be flown on ESA’s geostationary communication satellite Artemis, alongside the RIT-10 thruster, and usedfor north-south station keeping. A collaboration agreement, covering further development and marketing ofbigger diameter gridded ion thrusters for use on future large communication satellites, was signed between


Spacecraft CompositeMMS and DERA in 1997 as a response to the market assessments. Development of a 22 cm beam diameter ionengine T6 IPS with a nominal thrust of 150 mN is now in progress. This thruster is the basis for the Mercurymission.Radiofrequency Ion Thruster, 220 mm discharge chamber diameterThe RF ion thruster (RIT) principle has been developed at the University Giessen by Prof. Loeb. At Dasa(Daimler-Chrysler Aerospace) RIT thrusters are under development since many years. Discharge chamberdiameters from 10 to 35 cm have been investigated. The most advanced thruster system is based on a 10 cmdischarge chamber diameter (RIT 10). A thruster of this type has been flown on the retrievable carrier EURECAin 1992/93. Two thruster systems are currently installed on Artemis, where they will perform north-southstationkeeping together with the UK-10 thruster. In addition, function tests have been performed with the RIT35 (35 cm discharge chamber diameter), which can produce a thrust of 200 mN at specific impulses above 4000s. The ESA-XX has been developed in co-operation with AEA Technology in Culham (England) andLaben/PROEL in Florence (Italy), using the RF-ionisation principle of the RIT, the grid system design fromUK-25 and the neutraliser developed by Laben. The function tests of this thruster using a discharge chamberdiameter of 26 cm have been successfully completed in September 1998. Dasa has now started the developmentof a RIT ion thruster for commercial application with a nominal thrust level of 150 mN, which is expected to bequalified in 2001. The data used during this study are based on this thruster system.Table 6.2-2 shows main data, heritage and planned activities for the 3 electric propulsion systems.As an input to the mission analysis and trajectory calculations, the total power input (Figure 6.2-2) and totalspecific impulse (from the tank supply) versus thrust level have been estimated at Begin of Life (Figure 6.2-3)and End of Life (Figure 6.2-4) by Bassner (1999). The effect of lifetime on the thruster performance has beenconsidered as a reduction of specific impulse, which could be caused by erosion of the grid system for the ionthrusters and by erosion of the discharge chamber for the SPT thrusters. The necessary power input has beenkept constant. The operational data of RIT-XT and UK T6 are so close that additional mission analysis is notnecessary; therefore the diagrams are identical for RIT-XT and UK T6.RITA-XT SPT 140 UK-T6 IPSData source Dasa SEP MMSThrust level 50 - 200 mN 150 – 250 mN 40 – 240 mNThruster Dimensions 300 ∅ x 200 220 x 200 x 175 300 ∅ x 200Thruster Mass 6 - 7 kg 6.8 kg 5 kgFlowMassController2.7 kg (for 2) 0.6 kg (for 2) 0.7 kgPPU / PMA Mass 18 kg / -- 13.7 kg / ? 20 kg / 5 kgExpected Lifetime 10000 h 7200 h 10000 hHeritagePlanned activitiesRIT 10, 10 mN flown onEURECARIT 10 15 mN qualified forArtemisESA-XX tested at 200 mNRITA-XT: start ofcommercial development in1998, qualification expectedin 2001SPT-70 numerous flights inRussia;SPT 100, 80 mN flown onRussian Satellites, qualifiedfor western standardsThruster and electronicsunder development,qualification expected in1999Table 6.2-2: Electric propulsion system characteristic dataUK-10, 18 mN qualified forArtemisUK-25, 200 mN testedThruster and electronicsunder development,qualification expected in1999121


BepiColombo70006000RITA-XT, UK T6 1300 VRITA-XT, UK T6 1100 VRITA-XT, UK T6 900 VTotal Power Input, kW5000400030002000SPT, 300 VSPT, 400 VSPT, 500 VSPT, 600 V100000 50 100 150 200 250Thrust, mNFigure 6.2-2: Total power input versus thrust level.400035003000Total Isp BOL, s250020001500RITA-XT, UK T6 1300 VRITA-XT, UK T6 1100 VRITA-XT, UK T6 900 VSPT, 300 VSPT, 400 VSPT, 500 VSPT, 600 V10000 50 100 150 200 250Thrust, mNFigure 6.2-3: Total specific impulse, BOL.400035003000Isp EOL, s250020001500RITA-XT, UK T6 1300 VRITA-XT, UK T6 1100 VRITA-XT, UK T6 900 VSPT, 300 VSPT, 400 VSPT, 500 VSPT, 600 V10000 50 100 150 200 250Thrust, mNFigure 6.2-4: Total specific impulse, EOL.The operational data selected for this application are shown in Table 6.2-3.122


Spacecraft CompositeThrusters Total Power Consumption Specific Impulse BOL/EOLRIT/UK SPT RIT/UK SPTUbeam 1300 V Ubeam 600 V Ubeam 1300 V Ubeam 600 V4 x 150 mN 17400 W 14200 W 3500/3200 s 2200/1950 s3 x 200 mN 17340 W 14100 W 3550/3300 s 2300/2100 sTable 6.2-3: Operational data of the electric propulsion systemsIn the single-launch option, five thrusters are installed on the satellite as shown in the block diagram ofFigure 6.2-5. The mission can be performed with 4 ion thrusters operating in parallel at 150 mN (those in thefour corners of Figure 6.2-6). In case of a thruster failure the mission will be continued with 3 thrusters (on thediagonal of the square of Figure 6.2-6) at 200 mN. This allows to direct the thrust vector through the c.o.g. ofthe satellite in any case. The split-launch version has three thrusters on a line, used in turns (the central one at0.17 N in the early cruise, the outer pair at 0.34 N later on). Operation and failure tolerance of this configurationare discussed in Section 5.1.4.Figure 6.2-5 :Block diagram of the electric propulsion subsystem (single launch option).Figure 6.2-6: Thruster layout and redundancy concept (single launch option).All three thruster types fulfil the requirements, according to the investigations performed during the study.For the final application, a commercially available thruster system will be adopted. Adaptation of technicalparameters to the special requirements of the mission will be necessary and will have an impact on the costs.123


BepiColomboThe selection will mainly depend on the cost, which will again depend on the number of commercially producedthrusters and the necessary changes.6.2.3 Solar ArrayThe main requirements applying to the SEPM solar array are:• it shall provide the electric power required for the cruise to Mercury; the power demand shall be up to 10kW at 1 AU (depending on the thruster type used and the launch option);• it shall withstand the temperatures resulting from a Sun approach at 0.3 AU; temperature limitation shall beachieved with a 20% Second Surface Mirror (SSM) filling and a progressive tilting of the solar array awayfrom Sun, as the distance to the Sun decreases;• the solar array drive mechanism (SADM) shall provide rotation of ±180° around one axis perpendicular tothe trajectory plane (close to the ecliptic); no slip rings are required.There exists no detailed knowledge on the behaviour of solar cells at high intensities and high temperatures.Therefore the problem was approached with analytical modelling, based on supplier data as far as possible.Solar cell electrical characteristicsThe solar cells selected for trade-off were Silicon Hi-eta (16.3% AM0 efficiency) and GaAs/Ge (18%). Whenapproaching the Sun, both intensity and temperature increase. While increasing intensity increases also the solarcell power, increasing temperature reduces it. There exists a break-even point where the power gain by theintensity is compensated by the power loss due to the temperature. This point defines the maximum powerdelivered by the solar array. Further approach to the Sun reduces the power, until complete power loss due to avoltage breakdown when a specific temperature limit is reached. Keeping down solar cell temperature ispossible by application of optical surface reflectors (OSR), so-called second surface mirrors (SSM) and/or bytilting the array away from the sun. Both effects have been investigated. 10kW solar arrays have been definedand equipped with Si-Hi-eta and GaAs/Ge solar cells respectively. The Si solar array delivers its maximumpower at approximately 0.57 AU but needs to be tilted by at least 60° at 0.33 AU. Otherwise, the solar arraypower would break down completely. The maximum power is approximately 50% higher than the initial power,near Earth. The GaAs solar array has its maximum power at approximately 0.35 AU, close to Mercury'sperihelion. The solar array power never breaks down completely and the maximum power is approximately250% higher than near Earth. Therefore, a GaAs solar array has been selected for further investigations.In order to verify the assumptions made for the description of the solar cells at high intensities and hightemperatures, competent solar cell manufacturers were asked to provide experimental and theoretical data ontheir products in the intensity and temperature ranges considered. ENEL (Italy) provided a report evaluating testresults on single-junction and multi-junction GaAs-based solar cells (Campesato and Flores, 1998). For extremetemperatures and intensities a theoretical model developed by Dornier, and checked against a similar model ofENEL, was used.Mechanical LayoutThe solar array must have a minimum area of 55 m 2 in order to supply 10 kW at 1 AU. The ORION 2 solararray of the MMS EUROSTAR 2000 plus family meets this requirement. The 10 kW deployable solar arrayconsists of two identical wing assemblies. Each wing is made of one yoke and five panels. At launch, the wingsare stowed against opposite sides of the spacecraft. The panels are then deployed and locked The solar arrayincludes the following items:• solar panels, with substrates, solar cell circuits, and panel wiring;• yoke to support the panels and interface with the solar array drive mechanism (SADM);• stowage and deployment mechanisms;• wing wiring harness, insuring the connections between the panels at the yoke and SADM;• deployment indicators (microswitches) at each hinge line;• temperature sensors on the back surface of the outboard and centre inboard panels.The main dimensions of the 10 kW solar array wing with all hinges locked are shown in Figure 6.2-7.124


Spacecraft CompositeYOKEINBOARD PANEL CENTER INBOARD PANEL CENTER OUTBOARD PANEL OUTBOARD PANELYOKE HINGE2216(Panel)2250(Over All)1050(Hinge Line)CLOSED CABLE LOOP2788(Hinge Line)PANEL HINGE150271 Yoke plus 5 Panels2730(Panel)Figure 6.2-7: 10 kW SEPM solar array based on EUROSTAR 2000 plus (ORION 2).The 10 kW array described above is employed in the most power demanding option (Ariane 5 configurationwith ion thrusters). For the Ariane 5 configuration with SPT thrusters, the same array design was assumed, butwith four panels per wing (8 kW). For the split launch options, the array area was scaled down linearly (withdifferent panel sizes) to provide 5.5 kW at 1 AU.The basic panel substrate is a sandwich structure with perforated aluminum honeycomb core and highmodulus carbon fibre face sheets. The face sheets are filament-wound with local doublers and stiffening tapesadded to meet the natural frequency requirements with a minimum mass. An epoxy resin and curing process willbe selected to make the substrate tolerant against temperatures greater than 200 C. Each panel is 2730 mm inlength and 2216 mm in width. The total substrate thickness is 25 mm. Six hold-down points will be required tofix the folded wings to the spacecraft side wall at tolerable eigenfrequencies. The edges of the panels will bethermally protected with MLI superinsulation foil. A polyamide film (Kapton H) is bonded to the cell side facesheet, providing electrical insulation between the circuits and the conductive structure of the substrate. Twoextra layers of epoxy resin are used to bond the Kapton on the cured CFC-sheet. This prevents the carbon fibresfrom penetrating into the Kapton film and guarantees that no insulation failure will occur during all phases ofthe generator life. The yoke and hinges, the hold-down and release system, the deployment mechanism, and thesolar array drive mechanism are described by La Roche (1999).Electrical LayoutThe solar cell selected for the SEPM panels is a GaAs/Ge solar cell 41.0 mm in length and 42.4 mm in width.The panel size allows to connect 52 cells in series to form a string. Two strings in series form a circuit. Asection consists of 5 circuits having 104 solar cells in series and 5 cells in parallel. To improve thermal control,a row of second surface mirrors (SSM´s) with low absorptivity and high emissivity is laid down in the gapbetween each two adjacent strings. Each SSM has a size of 10 mm x 41 mm and is 100 µm thick.also arranged at the rim along the edges of the panels. The panel surface consists of 74.7% solar cells, 16%SSM´s and 5.7% substrate (gaps between cells and SSM´s). Instead of MLI-foil, SSM´s could also be appliedfor protection of the panel edges when they are fully illuminated by the Sun due to tilting of the array.Each panel is divided into 5 sections. Each section is magnetically compensated by the neighbouring sectionwhich has an inverted cell arrangement. Single sections are magnetically compensated by proper cable routing.Each section is individually connected to the power conditioning unit (PCU) which controls the operation pointon the IV-characteristic according to the power required by SEPM and the payload. The coverglass is a 100 µmthick ceria doped microsheet with enhanced anti-reflective or blue reflecting coating on the front side and IRreflectingcoating on the cell side. This reduces solar cell absorptivity by more than 10% and increases cellefficiency.Electrical PerformanceFor the reference array, the beginning-of-mission (BOM) power after deployment in a near Earth orbit will bein the order of 10575 W under the conditions assumed (autumn equinox, 90° sun incidence). The solar celltemperature will be around 25 C. Both power and temperature increase to the same extent as SEPM approachesthe sun. For an untilted solar array the temperature would increase to nearly 300 C at the end of the cruise phase(Mercury perihelion). Therefore, temperature limitation by tilting is required. Figure 6.2-8 gives the tilt anglesrequired to limit the temperature of the solar array to 150 C and 200 C, respectively. Figure 6.2-9 gives thepower as a function of Sun distance if the array temperature is limited to 150 C and 200 C, respectively.125


BepiColomboSolar Array Tilt Angle [°]70605040302010Tilt angle for T=150°CTilt angle for T=200°CFigure 6.2-8: Requiredtilt angle to limit solararray temperature to 150 Cand 200 C.00.25 0.37 0.50 0.62 0.75 0.87 1.00Sun Distance [AU]40,00035,00030,00025,00020,00015,000Power at T=150°CPower at T=200°CFigure 6.2-9: Solar arraypower vs. Sun distance withtemperature limited to 150C and 200 C.10,0005.000,00,00.250 0.375 0.500 0.625 0.750 0.875 1.000Sun Distance [AU]6.3 CHEMICAL PROPULSION MODULEThe main function of the Chemical Propulsion Module (CPM) is to host the bi-propellant propulsion systememployed for attitude control during the cruise, and for Mercury orbit insertion. The attitude control functionsare served by a redundant set of eight 20 N thrusters, while the planetary capture and Mercury orbit acquisitionmanoeuvres are performed by a 400 N engine. CPM also serves as a structural interface between SEPM and thescientific modules, and as support for additional equipment (split launch 1).The general features of CPM are common to all mission options. To realise a structurally efficientconfiguration, CPM is shaped as a short truncated cone, capping the SEPM thrust cone (Figure 6.3-1). Thepropellant tanks are supported by the main structure, and protrude below it into SEPM, in order to minimise thelength of the structure. The lower side of CPM is thermally insulated all over in order to provide thermalprotection after the jettisoning of SEPM.Figure 6.3-1: CPM configuration concept.In the mission options including the Planetary Orbiter (Ariane and split launch 2), CPM is mounted lateraly,in order to maintain the same attitude with respect to the Sun and planets as in the MPO operational phase, asshown in Figure 6.3-2. The separation interface with the Orbiter is made in three points, to minimise thermalleaks into the Orbiter after separation.126


Spacecraft CompositeFigure 6.3-2: CPM attached to Planetary Orbiter after SEPM jettisoning (split launch 2).CPM is smaller in the split launch configurations than in the Ariane configuration. The design is howeverstandardised to serve both payloads; only the upper module interface structure and propellant tanks differbetween the first and the second scnenarios. The consequence is that the design is inefficient in terms of massfor launch 1. As remarked in Section 6.1.2, this is an area for future work. Furthermore, CPM requiresadditional components (solar array and star sensors) for the cruise and the early Mercury orbit phase in launch 1(Figure 6.3-3).Table 6.3-1 shows the component list and mass estimate.Figure 6.3-3: CPM configuration for split launch 1.127


BepiColomboSingleLaunchMass (kg)Split launch 1(MMO+MSE)Mass (kg)Split launch 2(MPO)Mass (kg)Propulsion subsystem 69.8 33.8 33.8Thermal control 11.7 5.9 5.9Structure and mechanisms 78.0 29.6 29.6Interface to upper module 7.9 13.4 7.9Auxiliary equipment 0.0 6.7 0.0CPM Dry 167.4 89.4 77.1CPM Propellant 335.9 119.9 158.3Total wet mass 503.3 209.3 235.4Table 6.3-1: Chemical Propulsion Module mass budget. Mass items include margins according to itemmaturity6.4 MERCURY PLANETARY ORBITER6.4.1 ConfigurationThe configuration of the Planetary Orbiter is driven by the thermal design, the purpose of which is to reject asmuch as possible the very large heat inputs from the Sun and the planet, by means of high- efficiency insulationover the all body and a large radiator. The external shape (Figure 6.4-1 to 6.4-3) is a flat prism with slantingsides, tilted by 20° to reduce the view factor to the planet. Three sides (±X, +Y) are partially covered with solarcells, mounted on an Al substrate with, on average, a 30% cell filling factor, the remaining 70% being coveredwith Optical Solar Reflectors. Power is generated at any permitted solar incidence, due to the 20° slant.The –Y side is covered by a 1.5 m² radiator, the size of which is driven by the internal power dissipation(limited to about 200W); the relationship between dissipated power and radiator size is approximately linear.The radiator is protected from the Sun by rotating the spacecraft around the Z-axis by 180° every half Mercuryyear, and from the planet by a deployable shield, large enough to block the IR radiation for any permitted viewfactor to the planet (about 3.4 m² for 400 km minimum altitude). The IR shield is stowed at launch against theradiator, because of launcher fairing constraints, as shown in Figure 6.4-4.The externally mounted elements include:• three star sensors on the radiator side, protected from the Sun and planet radiation, and providing attitudemeasurement under all orbit conditions;• three Sun sensors for the safe mode, with a special HT design;• Nitrogen thrusters for reaction wheel unloading, in a balanced configuration;• UHF dipole antenna on nadir side (Lander);• three X-band low gain antennas;• deployable, 2-axis articulated, 1.5 m high gain antenna (HGA), mounted on a short boom close to theradiator side; a suitable latch mechanism restrains the antenna against the Orbiter -Z side at launch.The equipment (Figure 6.4-2) is mounted on the internal side of the +Z (nadir) platform, on a specialequipment platform at right angles, and directly on the radiator (including the battery). The equipment platformsupports the payload cameras and star sensors, to minimise the relative motion of the lines of sight. Instrumentapertures are protected by baffles and, when possible, by filters and MLI sheets (large FOV instruments). Forthe visual instruments, a special configuration is adopted, where the telescopes are placed at right angles tonadir, pointing towards dichroic mirrors (Section 6.4.2).A heat pipe network (embedded in the panels) connects the nadir and instruments platforms to the radiator.Another network connects the solar and Zenith panels to dissipate heat uniformly for different Sun illumination.The Orbiter structures are made of carbon-fibre reinforced plastics, in accordance with the radiationprotection guidelines (Section 4.2.2).The interface with CPM is reduced to three points on the -X side (to limit the heat input into the Orbiter afterseparation). This configuration permits, in principle, scientific observations during the cruise (e.g., Venus andMercury fly-by’s); however, attitude, thermal and communications constraints will limit such observations.Table 6.4-1 gives the Planetary Orbiter mass budget.128


Spacecraft CompositeFigure 6.4-1: Planetary Orbiter configuration.Figure 6.4-2: MPO configuration with solar panels removed, showing equipment layout.129


BepiColomboFigure 6.4-3: MPO dimensions.Figure 6.4-4: Planetary Orbiter launch configuration, showing IR shield and antenna stowage concept.Orbiter Mass BudgetMass [kg]Payload 60.0Data Handling incl. 2 µRTUs + 6 AOCS I/F boards 26.0RF 33.0Power & Harness 48.6Solar panel electrical components (panels included in structure) 5.6GNC/AOCS 23.6RCS incl. Propellant 24.7Thermal Control 47.1Structure and Mechanisms 88.7Total 357.3Table 6.4-1: Planetary Orbiter mass budget. Items include subsystem margins according to maturity130


Spacecraft Composite6.4.2 Thermal controlThermal designThe main features of the thermal configuration and thermal mathematical model (Rapetti 1999) are describedbelow and shown in Figures 6.4-5 and 6.4-6.A radiator is mounted on the -Y side of the spacecraft and is protected from the Sun by proper attitude, andfrom the planet heat flux by the IR shield. The external side is covered with Optical Solar Reflectors (OSR,α=0.2, ε=0.8 at end of life). The radiator area depends on the internal dissipation and the selected orbit. Therequired radiator area is about 0.7 m 2 for each 100 W dissipated.The IR shield is at right angles to the radiator. During part of the orbit, its –Z face is illuminated and somesunlight is diffused on the radiator. For this reason, the – Z side of the shield (toward the radiator) is coveredwith OSR’s in order to minimise the temperature and the diffuse reflection of solar radiation onto the radiator.Experimental measurements performed at ESTEC show that the diffuse component of the OSR reflectedradiation is less than 1%, thus reflection is basically specular and, due to the shield geometry, cannot impinge onthe radiator. On the planet side (+Z), IR MLI with Titanium foils is used due to the peaks of temperature beforeand after eclipses.The body panels on ±X and +Y sides, are tilted 20° to limit IR input; 30% is covered with solar cell stringsand 70% with OSR (average of the three panels). The panels are thermally separated from the S/C structure bystand-offs and MLI. Heat pipes are used to link the solar panels with the –Z side and to spread heat. OSR aremounted on the –Z panel that acts as an additional radiator and is thermally decoupled from the S/C with MLIand stand-offs.The HGA is made of an HT resistant composite with C-9 coating (α = 0.27, ε=0.7); C-9 is a rigid white glasscoating for flexible blanket insulation materials and is used on the Shuttle; it is composed of silica powder andcolloidal silica particles. The antenna boom and mechanism are wrapped with MLI (α = 0.2, ε=0.8) andconductively linked to each other.Most of the internal equipment and experiments are installed on the +Z panel. All internal items and thespacecraft walls are painted black with α = 0.9 and ε=0.9 to create a uniform internal environment. Each itemhas been modelled with one diffusive node connected with the other items and internal surfaces of the S/C byradiative conductors. Items mounted on the mounting plate and on the internal surface toward the nadir areconnected with these panels with linear conductors which take into account the filler between item and panel.The internal dissipation of the items has been considered constant, according to the power budget, for both thesunlight and eclipse phase; for the battery only a dissipation variable with temperature during eclipse has beenconsidered. The battery is mounted in a cavity insulated by MLI (α = 0.9, ε=0.9) from the main internalenvironment and by stand-offs from the internal surface toward the planet. The battery is connected by heatpipes to a dedicated radiator, taking up about 14% of the total radiator area. Heat pipes are embedded in thehoneycomb structure toward the planet and in the radiator; they are also used to transfer heat from instrumentsto radiator.On the experiments facing nadir, the external fluxes (solar, albedo, IR) are mitigated by filters or protections,as appropriate to each experiment. NAC, WAC, ALI, IMS, MGS units have baffles (α=0.9, ε=0.9), while MNSand MXS have a wall facing the external environment covered with MLI (α = 0.3, ε=0.05). The followingapproach is adopted for protection from thermal IR inputs through telescope apertures (Figure 6.4-7). Thetelescopes are mounted at right angles to nadir, and view a dichroic mirror in reflection; the wavelength banduseful for science is channelled into the telescope without the IR component, while thermal IR radiation istransmitted to the surface behind the dichroic and is cooled by radiation to space.Some instruments have low working temperatures (NAC< 90°C; IMS< 27°C; MGS


BepiColomboRadiator• OSR• honeycomb• bolted heat pipes (toconnect it with Zpanel)MLI under solarpanelsAdditional equipmentpanel:• honeycomb• bolted heat pipes(to connect itwith Z panel)IR shield:• honeycomb• OSR side –Z• MLI high temp.side +ZZ Panel for equipmentand experimentsmounting:• honeycomb withembedded heatpipes• ammonia constantconductance heatpipes• MLI high temp.side +ZFigure 6.4-5: Orbiter thermal design overview.Solar panel radiator:• 100% OSR,• Honeycomb high temp.• heat pipes hightemperature, embeddedPanel +Y• 70% OSR• 30% cells• Honeycomb high temp.• heat pipes high temp.,bolted, fillerPanel +X• 70% OSR• 30% cells• Honeycomb high temp.• heat pipes hightemperature, bolted, fillerStand-offMLIPanel –X• 70% OSR• 30% cells• Honeycomb high temp.• heat pipes hightemperature, bolted, fillerFigure 6.4-6: Orbiter solar panel thermal design overview.132


Spacecraft CompositeFigure 6.4-7: Approach to telescope thermalprotection from IR flux.Figure 6.4-8: Approach to MGS thermaldesign.Thermal AnalysisThe orbit considered in the analysis has 90° inclination, apocentre = 1500 km, pericentre = 400 km atsubsolar point at Mercury aphelion. The planet albedo is taken constant (0.11) and the solar constant iscalculated according to the distance from the Sun (6236 to 14358 W/m 2 ).The mean temperatures in the Orbiter have a seasonal behaviour, depending on the longitude of the Sun withrespect to the orbit plane, here designated by Ω (which is also associated with a variable distance to the Sun).With the orbital reference assumed in the analysis, Ω = 0° corresponds to the Mercury aphelion (subsolar pointon the side of the pericentre); Ω = 180° corresponds to the Mercury perihelion (subsolar point on the side of theapocentre); Ω =90° is the intermediate case with the Sun perpendicular to the orbit plane. Figure 6.4-9 shows themean internal temperature as function of Ω, assuming a constant power dissipation of 200 W and thermo-opticalproperties EOL. The figure shows that two extreme cases can be considered: Ω =90° (worst cold case) and Ω=180° (worst hot case).403020100-10-200 60 90 135 160 180OMEGA [deg]Figure 6.4-9: Mean internal temperatures as a function of sun longitude w.r.t. the orbit plane.More detailed system analyses performed during the study lead to the following conclusions. Temperaturesof the internal equipment are below 40 C generally and below 50 C in the worst case (perihelion). This confirmsthe results shown in Figure 6.4-9, which were obtained with a simplified thermal model. No layout or designoptimisation at single equipment item level was performed; an improved modelling is possible andrecommended for a future study. Externally mounted equipment (HGA, UHF antenna, solar sensors and solarmonitors) reach very high temperatures, up to 200 C, making ad-hoc designs necessary (recommended for thetechnology programme).For the experiments, special provisions must apply. The experiment apertures lead to very high IR fluxesreaching the interior. For the analysis, a simplified approach was adopted, simulating a degree of filtering of thefluxes. Further analyses are necessary in order to study in detail the design of each instrument and filters orprotections from the extreme external environment.Some specific analyses were carried out to verify individual instrument temperature requirements.133


BepiColombo1. Low temperatures must be provided locally for some sensors, by means of coolers with suitable heat sinks.The preliminary study performed for one experiment (MGS) shows the feasibility of coolerimplementation.2. The Mercury libration experiment imposes stringent requirements on the stability of the alignment betweenthe Wide Angle Camera and the star trackers lines of sight (order of 2 arcseconds on a time scale of 1orbit). A thermoelastic analysis showed that the maximum expected angular displacements are indeed ofthis magnitude, provided that the instruments are suitably isolated from the thermal deformations of themain structure. The effects of deformations internal to the instruments remain to be studied.3. The temperature excursion at the accelerometer level, on time scale of the order of the orbital period andlonger, must be


Spacecraft CompositeRF LinksRF LinksejectejectOrbiterRF COMS S/SHard LanderCOMS S/SAOCS/RCS AvionicsChemicalOrbiterPropulsionAOCS/RCSModuleRCSElectricalPropulsionModuleRCSejectejectC&DH-AvionicsCommand & DataHandlingT/C S/SejectSolarArraysPart of ejectableElectricalPropulsion ModulePower S/SPart of ejectableElectricalPropulsion Module ejectOrbiterExperimentsMag. Orbiter & Hard Lander I/Fduring transportationMag-SATreleasedfrom ChemicalProp. ModuleHard Landerreleasedfrom OrbiterFigure 6.4-10: Electrical configuration survey of Mercury Orbiter with propulsion modules interfaces(Ariane 5 launch option).TM V ideo /Di gital T MT C Vide o /D igita l T CHG A Poi ntingCon tro l (2 DO F)C&DHCommand &Data H andlingC D M SCon tr ol & D a ta M ana gem entSu ba ss em blyR e con fig .M od u le 1S a f e g u a rdM e m o r y 1C SS sAO C S Se nsor sMH SS TR sIM U(C& DH task in g& AO CS task in g)c o a r seS u nS e ns o rsMe r c u ryI n e r tia l3 S ta rHo r iz o nM e as u r e m .T r a ck e r sS e n so rUn i tHP C-1HP C-2S S MM10 G bitsm odul arM emoryBu s R TLin k I/FC PD UCL C WTra ns fer Fra m eG e ne ra to rPack et TCDec oderT C M A P1T im e &P ara me ter sT XO U SO 1 2T ime T ime r 1 r 2Ma ster C lo ckP r o c e s s o r W a t c h D o gP ro ces sor M od u leE RC 32& M e m o ryD AT A Bu sCo n t ro ll e rB CLo c a l B usS e ri a lL in k I /F&B a c k p la ne B usS erialLin kI/FA IU AO CS In te rfa ce Unit G imbals I/FMAC SBu sWD E1 - 44 RC Wh eelsR C S E le ct r. I/FEl . Pro p u ls io n I/FR CS El e ctr. I/FM PPTSt atusM o n ito rs fo r Te s tTest BusMonito rH igh R at e Data Acqui sition IEEE 135 5T BD serial Dat a Bu s, e.g. MIL 1553BRTUI/O C ha nn e lst o SA D rive El ectronicsA O C S P r o c e s s in gL in ks I EE E 1 3 5 5M a ste r Cl oc kL eg en d:: ho t re du nda nc ies: co ld red un dan tS y n c h . , C l o c k & T im in gPul se CM Ds, Status Mo ni to rin g& Th erm istor Acqu isitionMemory Lo ad Cmd s & 16 Bit Serial TMFigure 6.4-11: Avionics subsystem.Thermal control electrical items are the heaters and electrically controlled roller blinds (MagnetosphericOrbiter), thermistors, T/C power outlets in the PDU and thermostats for survival heater powering (especiallynecessary during the cruise phase). Nominally the temperature control is performed via S/W controlled circuits.Dedicated external functional/electrical interface to the S/C bus equipment, the experiments, theMagnetospheric Satellite and Lander check-out package, and the VMC camera are controlled from the C&DH.The MIL-STD-1553B is a promising candidate for a serial data bus system. Remote terminals of the data buswill be applied at users side as µRTU's if it is available in the design phase. A limited number of standard RTUchannels (TTC-B-01) as serial lines, pulse commands, and analogue & digital inputs will certainly be necessaryaccording to recurring units and for autonomous switching capabilities for configurations.To perform autonomous spacecraft control (Kahl 1999), the C&DH must have the necessary primary powerswitching capabilities to allow for autonomous S/C control and (re-)configuration and the AOCS must controlthe sensors and actuators. The switching capabilities can be implemented for the C&DH via the PDU and forAOCS either in the AOCS interface unit (ROSETTA approach), or via the same PDU interface as applied forthe C&DH. The C&DH and AOCS softwares and the reconfiguration module only need access to the PDU135


BepiColombopower outlets device driver (proposed baseline). Another major item for on-board autonomy is the classical hotredundancy of the COMS receivers and the packet TC decoders. These are powered by ‘Permanent Power’outlets from the PDU.t o/f ro m A O C S I/FSA Driv eC ontrolE lectronics• • •O rb ite r SAISA30V - 50VMainbusMain 28 VRegulator2 out of 3 MEAMPPTSA UMainbusErrorcold red.Amplifier50V - 100VPCULaunch PadPowerChargeReg. 1Battery 1Disch.Reg. 12 out of 3ChargeReg. 2MBBattery 2FilterDisch.Reg. 2Bus RTto/from RTUI/OLCLFCLLCLPyroModulePDUOrbiterPowerS/SSwitched PowerPermanent PowerSwitched T/C PowerPyro Circuitsto/from RTUSADMSA DMI SAU SAI/OPower SwitchElectric PropulsionPower (< 50V )MPP Status• ••MPPT2 out of 3Power UnitC ruiser SA 1• • •Cru is er SA 2Electric PropulsionPower ModuleFigure 6.4-12: Electrical Power Subsystem baseline.Planetary Orbiter solar arrayThe Orbiter solar array is body-mounted with 20° slant Kapton insulated panels on ±X side and +Y side(Figure 6.4-13). The area of the side panels is 1 m 2 , the area of the front panel is 1.7m 2 . Temperature ispassively controlled by the radiator (connected via heat pipes; heat conductivity 100 W/m 2·K) and secondsurface mirrors (SSM´s) on the panels (α = 0.045-0.09; ε = 0.81).+y+x+z136Figure 6.4-13: Orbiter solar array configuration.The same materials are foreseen as for the SEPM solar array with the exception of the solar panel substrates,which shall be either 2mm thick Al-plates with 50µm thick Kapton insulation foil, or CFC-honeycomb withintegrated heatpipes. The latter concept, which has a mass saving potential of more than 20 kg, is the referencedesign adopted for this study. The solar cell selected for the Orbiter panels is a GaAs/Ge solar cell of the sametype as for the SEPM solar array, but only 20 mm in length and 40 mm in width. Sixty solar cells are connectedin series to form a circuit. The side panels are equipped with 10 circuits each, the front panel with 8; the fractionof panel area covered with cells is 44% (X-panels) and 20% (Y-panel). Second surface mirrors (SSM´s) withlow absorptivity and high emissivity cover the remaining area of the panels. The circuits of each panel areconnected in parallel and commonly controlled by the power conditioning unit which tracks the operation pointon the IV-characteristic according to the power required by the Orbiter and the payload. Therefore, the solararray generates only as much electrical power as it is necessary to serve the load and charge the batteries (plusefficiency losses from the PCU/PDU). No electrical excess power is generated due to high insulation, e.g. inMercury perihelion, instead higher insolation will be reflected and absorbed according to the thermal


Spacecraft Compositeabsorptivity and reflectivity of the solar cell characteristics. The coverglass shall be a 100 µm thick ceria dopedmicrosheet with enhanced anti-reflective or blue reflecting coating on the front side and IR-reflecting coating onthe cell side. This reduces solar cell absorptivity by more than 10% and increases cell efficiency.The +Z-axis of the Orbiter is always nadir pointing. Therefore, depending on the ascending node, the solaraspect angle of the solar panels changes continuously. Limit Sun aspects are reached by the solar array onterminator orbits (constant power provided by the Y-panel) and subsolar orbits (power provided by the ±Xpanels, cycled with eclipse). Moreover, the array power output depends on the distance from the Sun, withlimiting cases represented by aphelion and perihelion. The maximum solar array temperature in perihelionposition is 193 C for the subsolar orbit (no power consumption assumed). The orbit-average power delivered inthe nominal 400 x 1500 km polar orbit is shown in Figure 6.4-14. Limiting values are 390 W at aphelion, 550 Win the terminator orbit, and 680 W at perihelion.12001050Array power (W)90075060045030015000 4 8 12 16 20 24 28 32 36 40 44Time from perihelion (days)Figure 6.4-14: Orbit-averaged power delivered by the Planetary Orbiter solar array. The dashed line isthe power delivered by the Y-panel, the dotted line is the power delivered by the X-panels.Power [W] Dissipation [W]Sunlit Eclipse Sunlit Eclipsea. Perihelion Payload 89.4 85. 67.9 63.5RF 77. 77. 51.9 51.9Data Handling 19.8 19.8 19.8 19.8Power 5.5 5.5 15.5 15.5AOCS/GNC 28.6 28.6 28.6 28.6Total load power 220.3 215.9 183.7 179.3System margin (10%) 22. 21.6 18.4 17.9PCDU and harness loss (9%) 24. 23.5 0. 0.Total power demand 266.3 261. 202.1 197.2Eclipse energy (Wh) 134.7Array power 419.6b. Aphelion Payload 60.8 60.8 39.3 39.3RF 61.1 61.1 42.3 42.3Data Handling 19.8 19.8 19.8 19.8Power 5.5 5.5 15.5 15.5AOCS/GNC 28.6 28.6 28.6 28.6Total load power 175.8 175.8 145.6 145.6System margin (10%) 17.6 17.6 14.6 14.6PCDU and harness loss (9%) 19.1 19.1 0. 0.Total power demand 212.5 212.5 160.1 160.1Eclipse energy (Wh) 146.8Array power 389.4Table 6.4-2: MPO operational power budgets and power dissipation budgets. Subsystem powerrequirements include 10% margins137


BepiColomboThe power demand on the Orbiter must be kept under strict control because of (1) the power dissipation limitand (2) the solar array power limit. The ∼200 W dissipation limit arises because of thermal reasons; the power islimited by the size of the array that can be accommodated without steerable wings and the required fraction ofoptical solar reflectors. In the perihelion case, constraint (1) applies. The power budget (Table 6.4-2) shows thatthe dissipation is reasonably close to the goal even in the unlikely case (assumed in the table) of operation of allpayload items at the same time. In the aphelion case, constraint (2) applies, and a power cap of 60 W was put onthe payload items (remote sensing + radio science). Notice that the power shortage at aphelion is of shortduration, and the power needed for full payload operation becomes available ~5 days before or after aphelion.Experiment time sharing is anyway needed because of the telemetry limits, as discussed in Section 5.4.1. Theobservation plan of the Orbiter will have to take into account the thermal/power constraints as well as thetelemetry constraints.6.4.4 CommunicationsThe telecommunication elements in the Mercury spacecraft comprise:• Earth telecommunication elements (X/Ka-band), to provide the means for tracking, telemetry and commandin all mission phases; the Earth TLC elements reside in MPO, both during the cruise, when the Orbiter is apart of the composite, and during the Mercury orbit phase;• data relay telecommunication terminals (UHF band) on the Orbiter and MSE.Earth telecommunicationsTelecommunications to Earth are constrained by the variable distance to Earth (both during the cruise and inorbit), by occultations and by radio blackouts. On a typical mission profile, the minimum distance in orbit is0.57 AU (85 x 10 6 km) and the maximum distance is 1.43 AU (214 x 10 6 km); the ratio of maximum tominimum distance is about 2.6 and the time from a maximum to the next minimum is about 58 days (half asynodic year). Therefore, ∼6.5 times more data can be transmitted at closest approach than at maximumdistance; hence, the data return strategy must be planned over a synodic year of 116 days.Telecommunication with the Earth takes place when the ground station is available, and in absence of radioblackouts or occultations. Radio blackouts occur when Mercury is aligned with the Earth and Sun, so that theground antenna sees the spacecraft with the Sun either in the background or in the foreground. This conditionoccurs regularly, every 58 days. The angular separation between Mercury (the spacecraft) and the Sun takesminimum values that depend on the epoch, as shown in Figure 6.4-15. The minimum angles fortelecommunications are estimated to be ≈ 2.8° (Sun between Mercury and Earth), and 0.33° (Sun behindMercury). On a yearly average, blackouts subtract about 5% communications time. Occultations occur when theEarth is close to the plane of the spacecraft orbit around Mercury. The selected antenna articulation conceptincreases the occultation time; the average occultation per year then adds up ~20%.Assuming that the Perth 35m dish is available for 8 hours per day with a spacecraft elevation > 10°, and aftersubtraction of occultations and blackouts, it appears that 25% of the time is available for communications in the400 x 1500 km orbit.3025sun-earth-mercury angle (deg)20151050800 880 960 1040 1120 1200time (days)Figure 6.4-15: Example of evolution of the Sun – Earth - spacecraft angle.Earth telecommunications subsystem. The telecommunications subsystem shall establish and maintain thetelecommand and telemetry links with the Earth station during all mission phases. The telemetry rate wasdetermined during the study, as a compromise between scientific needs and engineering constraints. Twooptions are considered:138


Spacecraft Composite• X/X system (X up – X down)• X/Ka system (X up/down – Ka down).The final choice is in favour X/Ka band system which has a higher-capability.Figure 6.4-16 shows a block diagram of the X/Ka-band system. The proposed on board system architectureincludes:• redundant X-band chain for telecommand reception and low rate telemetry transmission (cruise and onorbitemergency mode);• redundant Ka-band chain for high rate science telemetry transmission (Mercury orbit);• X/Ka-band transponder;• 20 W RF output power HPA’s;• 1.5 m dual-band (X/Ka) High Gain Antenna reflector, 39.5 dB peak gain, 2-axis steerable over a large solidangle;• 3 Low Gain Antennas (X band) with 9.5 dB boresight gain, 0 dB at 50°.The architecture allows sharing high power amplifiers for simultaneous telemetry downlink and radio sciencesignal transmission. The reference ground station is Perth, upgraded to X-band receive / transmit capability andKa-band receive capability. Ka-band uplink is needed by the radio science experiment, and the correspondingground station equipment will be provided as a PI supplied item.Operating modes and telemetry rates. Four high level telecommunications modes are envisaged:• LEO mode, used in the LEOP phase;• Deep Space HGA mode 1: a suppressed carrier mode used, when ranging is not required, to transmit highdata rates not allowed by the CCSDS standards in residual carrier mode. Only Doppler tracking is availablein this mode;• Deep Space HGA mode 2: a residual carrier mode, allowing ranging as well as Doppler, with reduced datarates;• Deep Space LGA mode: used in case of on board failures that can affect the attitude (transition commandedby the on board control computer).Mode transitions are commanded by the Earth station via immediate or time tagged commands, except in theLGA mode where it is automatic. Table 6.4-3 shows the telemetry rates available in each mode in LEO and atthe maximum and minimum spacecraft distances from Earth during the scientific mission. The Deep SpaceLGA mode is an emergency mode and a 34 m DSN antenna is assumed (better EIRP and G/T) to increase thedata rate, which would otherwise be very low at the edge of the LGA beamwidth. Ka-band telemetry with theHGA increases the data rate by a factor of 4 to 5 with respect to X-band. The suppressed carrier mode allows asignificant increase of the data rate (a factor of 2 at X-band, 1.5 at Ka-band). The residual carrier mode is usedduring the range measurements only (2-minute sessions every 15 minutes).During the mission, the telemetry rates must be varied in discrete steps to cope with the varying distance ofthe spacecraft. Figure 6.4-17 shows the telemetry rate as function of time at Ka-band (32 GHz), assuming a 20W RF power and a 1.5 m antenna reflector, in the suppressed carrier mode.Data return of the scientific mission. The data returned to Earth by the telecommunications system in amission lasting 1 year can be calculated from the average telemetry rates, after subtracting the periods when thelink is not achieved because of ground station availability, occultations and blackouts. The nominal telemetrymode is Deep Space Mode 1 (suppressed carrier) except in the periods when the ranging is on, with reduceddata rate. Taking into account the switching between modes, the estimated data return after 1 year at Ka-band is∼1550 Gb.High Gain Antenna configuration and design. The design drivers of the High Gain Antenna subsystem are theenvironment (high operation temperature and temperature X/Ka-band excursions, dimensional stability, pointingstability, radiation) and the RF design (dual frequency X/Ka-band, steering range larger than 2B ster, andlimited envelope for spacecraft accommodation: radiating aperture limited to 1.5 m, axial length less than 1 m).Three configuration options were traded-off: (1) single on-set reflector illuminated by dual-band feed; (2)Cassegrain on-set antenna with metallic subreflector illuminated by dual-band feed at secondary focus; (3)Cassegrain on-set antenna with dichroic subreflector, reflective at Ka-band and transparent at X-band.Configuration 3 was selected as a design reference (performance can be optimised at each band; minorblockage; straightforward feed design; configuration is flexible even in case of reduced aperture). It has a strongdesign heritage from CASSINI. Configuration 2 is retained as a possible alternative for its advantages in termsof mass, absence of transmission lines, thermal protection of the feed package. It has inferior performance at X-band, and it would be optimised for Ka-band. It would not be viable if the aperture were to be further reduced.139


BepiColomboThe proposed antenna (Figure 6.4-18) has a circular reflector diameter of 1.5 m, focal length of 0.63 m andF/D ratio of 0.42. With the above mentioned geometrical parameters, the peak directivity is D = 41.00 dBi, andthe peak gain after losses G = 38.25 dB. The feed aperture is about 60 mm and the overall length (Horn +Septum Polariser) about 180 mm. The feed is supported in the focal point by means of four struts converging atthe feed interface flange. A suitable backing structure is provided to properly support both main reflector shelland feed supporting structure. The feed supporting structure is made of four hollow struts connecting ribs edgesections toward the feed mechanical interface flange. Two wave guides WR112-RH (reduced height) will runfrom the feed polariser interface toward the reflector back side along the struts and radial ribs (structural supportto wave guide path). The selection of technology and materials is the subject of the antenna technologyprogramme. The mass estimate, based on a light-weight composite design, is 9.2 kg.Ka RadioScienceHigh Gain X- Kaband AntennaDiplexerBand Pass FilterLow gain X bandAntenna 1 and 2DiplexerDiplexerSW-2DiplexerSWSWSW -1K band TWTA K band TWTAX bandTWTAX bandTWTA3 dbPWR COMBCAMPPWR COMBCAMPPWR COMBCAMPPWR COMBCAMPX bandRadioScienceDownlinkTo DH TC DecoderFrom DH TMFormatterRadio ScienceKa TransponderX-BandReceiverRNGX-BandTransmitterK-BandTransmitterX-BandReceiverX-BandTransmitterK-BandTransmitterRNGToDHTCDec.FromDHTMForm.Uplink Carrier To RadioScience ExperimentUplink Carrier To RadioScience ExperimentFigure 6.4-16: X/Ka-telecommunication system block diagram.140


Spacecraft CompositeData rates at Ka BandData rate [bps]400,000350,000300,000250,000200,000150,000100,00050,0000811.164826.838849.230879.928921.303932.756Mission elapsed time [days]974.223991.3341029.7321046.3241079.1131091.9021141.9561165.355Figure 6.4-17: Telemetry rate as function of mission time (Ka band, suppressed carrier).Telecommand, X band, 7.18 GHzLEO 0.55 AU 1.48 AULEO mode 2 kbps -- --Deep Space HGA mode -- 2 kbps 2 kbpsDeep Space LGA mode (DSN 34m) -- 60 bps 20 bpsTelemetry, X band, 8.4 GHzLEO mode 4 kbps -- --Deep Space LGA mode (DSN 34m) -- 24 bps 4 bpsDeep Space HGA mode 1 (suppressed carrier) -- 72 kbps 12 kbpsDeep Space HGA mode 2 (residual carrier + ranging) -- 54 kbps 8.6 kbpsTelemetry, Ka band, 32 GHzDeep Space HGA mode 1 (suppressed carrier) -- 350 kbps 58 kbpsDeep Space HGA mode 2 (residual carrier + ranging) -- 270 kbps 36 kbpsTable 6.4-3: Operating modes and telemetry ratesFigure 6.4-18: Selected antenna geometry.141


BepiColomboThe Antenna Pointing Mechanism (APM) provides simultaneous, continuous rotation of the HGA in bothazimuth and elevation. The APM will consist of pointing mechanism with azimuth/elevation motor drive unitsand X/Ka-band rotary joints, APM electronics and supporting structure.The articulation concept is designed to provide Earth coverage for a large fraction of the orbit period, usingthe fact that the orbit is polar and the Earth direction remains close (within ±11°) to Mercury’s equator. Thereflector is placed on top of the Orbiter on a short boom; a mechanism at the boom root provides the rotation inazimuth, while another mechanism at the reflector rim provides the elevation degree of freedom.Communications are restricted to the orbit arc on the side of the Earth. Coverage is maximised by allowing theantenna to ‘nod’ towards the Orbiter roof. Figure 6.4-19 shows the azimuth and elevation degrees of freedom.54% to 100% of the orbit period is covered, depending on the longitude of the Earth w.r.t. the orbit plane. Theyearly average is 77%.The proposed technology development programme includes the 2-DOF antenna pointing mechanism androtary joints (under study within ESA TRP), high temperature materials (basic reflector materials, dielectricmaterials, metallic materials for RF devices, metallic and non-metallic materials for fasteners, inserts, tensioningelements, threaded bolts; antenna pointing mechanism and kinematic parts), and high temperature thermalhardware (paints/coatings, thermal blankets, protective shields).XoYoφ= -90° φ= 0° φ= +90°XoZoψ= 90° ψ= 0° ψ= -25°Figure 6.4-19: Antenna degrees of freedom in azimuth (φ) and elevation (ψ).The rotation angles arecounted from a reference position φ = 0°, ψ = 0° in which the antenna boresight is parallel to the X-axis.The azimuth spans up to -90° ≤ φ ≤+90° or +90° ≤ φ ≤+270° in a single slew (no full 360° turn required).The elevation spans -25° ≤ ψ ≤ +90° in a single slew. After an actuation cycle, the antenna must slew back- in azimuth only - to the position required for the next cycle.Data relay telecommunicationsData relay is implemented between the Planetary Orbiter and the Surface Element in the single launchscenario. Data are collected by the Surface Element at low rate and transmitted to the Orbiter at each overheadpass. In the single-launch scenario, an option was addressed in which the capability for inter Orbiter (MMO-MPO) communications exists as well. One of the conditions is that the orbit period of the MagnetosphericOrbiter be an integer multiple of the period of the Planetary Orbiter, so that the communication windows repeatat regular intervals. Although this options was later discarded in favour of independent Earth communicationswith the two Orbiters, the ‘resonant’ orbit design (T MMO = 4 x T MPO ) was maintained.The generic design requirements and assumptions were as follows:• UHF-band for minimum free space loss: Forward 437 MHz, λ=70 cm; Return 401 MHz, λ=75 cm;• CCSDS Short Range data relay standard applicable;• Minimum bit error rate: 10 -6 ;• BPSK Modulation with Reed-Solomon and Viterbi coding; required Eb/No=2.7 dB; Turbo coding as anoption;• wide antenna patterns with maximum gain in broadside direction;• Doppler ranging only.The design assumptions and link analysis results are summarized in Table 6.4-4.142


Spacecraft CompositeData rate / LifetimeLander output RF power 0.5WAntenna gains @ θ = 40°Limit slant rangeDuration of comms windowNumber of passes in 7 days 72Data return in 7 daysSurface Element to MPO4kbps / 7 daysOrbiter 2 dB, Lander -1 dB1000 km8 min138 MbitTable 6.4-4: Data relay assumptionsThe useful link window is identified at a slant range ≤ 1000 km, when each element has the other within 45°of its antenna pattern symmetry axis. The contact time lasts 8 minutes per orbit (Figure 6.4-20), and the data rateavailable from the link budgets is 4 kb/s with 0.5W RF power (including coding).9080Slant range (100km) & elevation (deg)70605040302010090 100 110 120 130 140Time from pericentre (min)Figure 6.4-20: Orbiter – Lander communications window. The continuous line is the slant range, inunits of 100km. The dashed line is the elevation above the lander horizon, in degrees. Dotted vertical linesidentify the permitted communication intervals.For the lander, a deployable turnstile antenna was preliminarily selected to sustain the high decelerations,with a radial wire grid as ground plane (Figure 6.4-21). This antenna can be constructed with titanium to sustainthe high environmental temperatures. Its mass is only 300 grams. The antenna is assembled in a so-called crossdroopingdipole configuration. Unlike the ordinary dipoles, the ends of the radiators are lowered ('drooping').Changing the inclinations of the dipole arms controls the radiation pattern of the antenna, while acting on thedipoles diameter varies the bandwidth. Moreover the drooping dipole antenna can be made into a multiband orvery wideband antenna by connecting several pairs of drooping dipoles in parallel.Figure 6.4-21: MSE Antenna layout.The link will be enabled on Orbiter commands. To save battery energy, the receiver could be enabled by asequencer some time before each predicted link interval. For the Planetary Orbiter antenna, a crossed droopingdipole fits the accommodation constraints, and can suffer the heavy thermal operating conditions with minorimpairment on bandwidth. The data relay assembly is made up of directive UHF antenna and UHF transceiver.143


BepiColomboThe transceiver includes Receiver/ Demodulator / Transmitter, integrating in a single box all RX and TXfunctions and the interface with the Data Handling subsystem. The transceiver dimensions are (mm) 140 x 106 x32 (commercial off-the-shelf) and the weigh is 1 kg. Two units are accommodated for redundancy.6.4.5 Guidance, Navigation and ControlThe driving requirements on the GNC/AOCS subsystem are (Martella 1999):• autonomy, because of the long distance to Earth (two way light travel time ranges from 9 minutes at closestapproach to 24 minutes at 1.5 AU) and limited ground station availability (8 hours per day);• reliability, because of the harsh environment (temperature, radiation);• accuracy, imposed by the payload camera observations in Mercury orbit.144Attitude disturbancesThe principal disturbing effects are due to (a) solar radiation pressure, both during the cruise and orbit, (b)planetary radiation pressure in Mercury orbit, and (c) disturbances induced by the electric and chemicalpropulsion during the cruise.For the scientific phase, computation of the radiation pressure torques was performed in the extreme cases ofperihelion and aphelion. The maximum torque about the X-axis is on the order of 4 x 10 -4 Nm (perihelion).Moreover, as the orbit around Mercury is eccentric (e=0.1623), a portion of angular momentum must beallocated to follow orbit rate variations; the maximum torque required is 10 -4 Nm. The total angular momentumto be compensated depends on whether thrusters (non-cyclic) or reaction wheels (cyclic) are used. The largefraction of cyclic torque permits saving about 60% propellant if the control is based on reaction wheels(reference design solution)During the cruise, the spacecraft composite is subjected to disturbance forces and torques due to solarradiation pressure. Figure 6.4-22 shows the torque and force profiles in a typical cruise case, including theeffects of time-varying distance from the Sun, the variation of the position of the centre of pressure and centre ofmass, and the area and optical characteristics of surfaces subjected to solar radiation. Peaks correspond to thepoints closest to the Sun (around 0.3 AU). These torques will be compensated by the CPM bipropellantthrusters. Some mass savings could be achieved by compensating the flux torque by small rotations of the solararrays used as solar sails.The electric thrust is produced by jets parallel to the longitudinal axis of the cruise composite. Anymisalignment will produce a torque which must be compensated. Misalignments can build up to 2°, as aconsequence of erosion processes inside the nozzle. Figure 6.4-23 (left panel) shows the resulting torque buildupin a worst-case analysis (single-launch option). The increasing trend is caused by both the linearmisalignment and the increase of moment arm associated with tank depletion. As the accumulated angularmomentum would be unacceptable, another control strategy for pitch and yaw was adopted, based onmodulating the thrust of each thruster (four in this example), to compensate the cross torque while maintainingthe nominal thrust constant (in the axial direction). Modulation up to ±30% of nominal thrust is required (Figure6.4-23 right panel), fully within the thruster range. This strategy produces spurious forces normal to thelongitudinal axis, which are of no concern since the whole spacecraft can be rotated through the small anglerequired to align the total thrust vector to the required direction. The roll torque (which is anyway the smallest)cannot be compensated in this way in case of one thruster failure, and therefore the associated momentum istaken into account in the propellant budget. Again, the CPM thrusters are used for torque compensation.Another source of disturbance is the torque due to misalignment of the 400 N engine of the ChemicalPropulsion Module during the insertion and orbit acquisition manoeuvres. This high disturbance torque is adesign driver since (a) to guarantee the needed control authority, a minimum thrust of 15 N is required (20 Nassumed for torque margin); (b) to rapidly counteract this torque and avoid destabilisation, the controller mustwork at a sufficiently high frequency (10 Hz taken as limiting threshold).GNC/AOCS architecture and sensorsThe approach selected for managing attitude determination both in cruise and on orbit is based on data fromtwo out of three non-orthogonal medium-FOV star sensors, supplying high accuracy estimation of attitude oneach axis; the information rate needed for control is derived from the attitude information via proper filtering.The implementation relies on the so-called Stellar Autonomous Attitude Determination System (SAADS), asystem developed by an industrial team guided by Alenia. In the multi-head scenario the SAADS software isable to combine output from two out of three optical heads to determine both attitude and rate (up to 5 deg/s), byapplying adaptive estimation techniques.


Spacecraft CompositeThe system outputs high accuracy attitude and attitude rate information in inertial coordinates, without anysupport from ground (apart from orbit updating around Mercury) or other onboard sensors, after onboardprocessing of the data provided by a set of optical heads, rigidly mounted on the spacecraft body. It implementsa fully autonomous gyroless system that can recover from the "lost in space" condition without any externalinitialisation. The SAADS includes two embedded compressed star catalogues ("image" catalogue, to recoverfrom "lost in space" conditions, and "reference star" catalogue, for the normal mode operation, to supportpointing and slews), image recognition and prediction correction algorithms. It can handle partial or totaloccultations, and reconfigure itself between different operating modes and is insensitive to extended objectsinside the frame.The SAADS software is executed in the CDMU ERC32 based computer. The SAADS software manages theprocessing of CCD frame as well as the attitude and rate determination; autonomous star trackers are thereforenot required and three simpler and cheaper optical heads are sufficient. The interface between the CDMUprocessor and the optical heads will be chosen on the basis of the computer and data handling system design. Apoint-to-point bi-directional serial link RS422 like with IEEE 1355 protocol for high rate communication at 25Mb/s could be suitable. SAADS is able to handle two active out of 3 OH’s by multiplexing in time commandsand data acquisition to the different sensors. Each OH is operated at its nominal working period (200 ms), butthe SAADS measurement filter is fed every 100 ms with new data coming from one of the OHs. In such a way,the attitude and rate estimates can be supplied in output up to a frequency of 10 Hz. Thus, the pointing accuracyand stability requested for nominal modes, together with the requested update rate for boost modes, are achievedthrough star data fusion and proper filtering techniques.20.6radiation pressure force (mN)1.510.5radiation pressure torque (mN m)0.40.200 200 400 600 800mission elapsed time (days)00 200 400 600 800mission elapsed time (days)Figure 6.4-22: Typical profiles of solar pressure force and torque during the cruise.0250Cross-axis torques (mNm)102030Thrust (mN)200150100400 500500 500 1000mission elapsed time (days)Mission elapsed time (days)a: uncontrolled torque b: thrust magnitude modulationFigure 6.4-23: Uncontrolled electrical torque build-up and required thrust magnitude modulation.The 3 OHs are mounted in a plane 30° skewed w.r.t the XY-plane to avoid the Mercury limb. Two axes arerotated ± 60° w.r.t. YZ-plane, the third lies in the YZ-plane, as shown in Figure 6.4-24.145


BepiColomboOH 160°OH 2SAS260°30 °60 °YZXOH 341°YZX41°60°SAS1SAS3146Figure 6.4-24: Star tracker optical heads, SAS and MHS layout.In addition to the star cameras, the onboard sensor complement includes linear accelerometers, ring lasergyros (RLG), analogue sun sensors (SAS) and a Mercury horizon sensor (MHS).A Linear Accelerometer is integrated inside the Inertial Measurement Unit (IMU) together with Ring LaserGyros. During main engine boost or trajectory correction manoeuvres it signals when the measured ∆v reachesthe prescribed value commanded by ground. RLG are used to damp rates at launcher separation and tomanoeuvre spacecraft into a safe attitude after a fallback; thereafter, they are actively used until nominaloperating modes are enabled from ground. Complete coverage of the sky is requested to permit single-failuretolerance in case of RLG partial failure. Due to the variable attitude of the thermal shield three SAS sensors arerequired (Figure 6.4-24). Medium accuracy (0.25-0.5°/s) should be sufficient to cope with the safe mode relaxedpointing requirement.The MHS is devoted to managing the safe mode (accuracy ≈ 1°) when orbiting Mercury; usage duringscientific observations to improve orbit parameter estimation could be evaluated too. The sensor, which isderived from a low altitude conical Earth sensor, is formed by two telescopes (FOV = 22°) mounted in the XYhalfplane, 41° skewed w.r.t. +Y and –Y (Figure 6.4-24), tailored to the prescribed altitude range (400-1500km). The combination of the measurements from the two telescopes permits the measurement of roll and pitchangles during the whole orbit. In safe mode, MHS outputs are used to correct attitude estimates in a gyrocompassingstrategy where the prediction is based on RLG measurements.Two emergency modes are envisaged, for the cruise phase and the scientific phase. The former is based onSAS and RLG; the latter relies on the MHS and RLG. The need for a MHS during scientific operations arisesfrom the orbit thermal constraints. Due to the high Mercury infrared flux, the satellite must maintain the nadirpointed attitude. As the mode must work also in case of prolonged ground link outage, the mentioned attitudecan only be guaranteed by providing the satellite with a sensor able to compute the roll and pitch angles in aMercury centred LVLH frame. The safe mode in the cruise scenario can be managed with a simpler approachwhere two axes are observed by using SAS measurements while the third is only observed in rate via RLGinformation.ActuatorsThrust during the cruise is supplied by electric thrusters. Depending on the mission option, five (singlelaunchoption) or three (split-launch option) are employed. The operation and redundancy concepts are inSection 6.2.2.Attitude control actuation during the cruise is supplied by eight 20 N bipropellant thrusters placed in theCPM. Such powerful jets are needed to compensate the worst-case disturbance when the 400 N boost isactivated; they are also convenient to perform faster manoeuvres (about 600 s for a 180° rotation). However, 20N thrusters could induce a quite large ripple in the control during the cruise. Therefore, during the cruise thedisturbance torque will be compensated by modulating the electrical thrusters themselves, while the use ofbipropellant thrusters will be limited to roll axis control. During the 400 N boost, the spacecraft control will bemaintained by using 4 x 20 N bipropellant thrusters.Simulations of Planetary Orbiter control were performed assuming cold-gas thrusters, reaction wheels andFEEP thrusters. Cold-gas based control (8 x 5 mN thrusters required to guarantee redundancy) has twoimportant drawbacks: the high propellant consumption (more than 30 kg) caused by the tight dead-band, andviolation of the acceleration requirements from the payload accelerometer. Wheel-based control permits a betterperformance and saves about 10 kg w.r.t. the first strategy; the acceleration requirement is violated during wheeloff-loading phases but the impact is negligible at the expected off-loading rate (maximum once per orbit). TheFEEP-based option is the best from the attitude determination (no gyroscopic coupling) and control viewpoint(lower short-term errors and jitter) and does not violate the acceleration requirement. The reaction controlsystem mass would be small (15 kg including caesium); the drawbacks are the high power demand (up to 100 W


Spacecraft Compositepeak) and contamination. Wheel based control was selected as a preliminary design baseline. A set of fourwheels able to supply a torque of about 10 -2 Nm and to charge an angular momentum of a few Nm; it guaranteeswide control margins, maximum off-loading rate of once per orbit, and manoeuvrability.The main features of the different thruster sets are summarised in Table 6.4-5. The propellant budget,including 50% margins, include 40 kg of bipropellant and 9 kg of cold gas.Operative modes and AOCS logicThe mode transition diagram is shown in Figure 6.4-25. After separation from launcher, Safe AcquisitionMode (SAM) is enabled to acquire the Sun. As STR optical heads are still occulted by the thermal shield, thisphase is managed using SAS and RLG measurements. When the Sun is acquired, solar arrays and thermal shieldare deployed; then the star sensor optical heads are available and Initial Attitude Mode (IAM), based on SAADSpattern recognition mode, is enabled to initialise attitude. Thereafter, Nominal Operative Mode (NOM) isactivated and the satellite is rotated to the prescribed orientation. Attitude is estimated via star sensors both inpointing and slew phases. During the cruise, control reference is computed via interpolation of Sun and Earthlocations contained in a time-tagged table; during Mercury observation, it is obtained by propagating orbitonboard. Orbital parameters are computed on ground, based on tracking and Doppler measurements.SOLAR ELECTRIC PROPULSION MODULEElectrical (Xenon)5 x 0.2 N (single launch)3 x 0.2 N (split launch)Cruise delta-VCHEMICAL PROPULSION MODULEChemical (bi-prop) 1 x 400 N Mercury insertionAnd apocentre reductionChemical (bi-prop) 8 x 20 N (2 branches) Spacecraft attitude controlduring cruise & insertionPLANETARY ORBITERChemical (cold-gas) 8 x 0.020 N Wheel off-loadingTable 6.4-5: Thrusters summaryWheel Off-loading Mode (WOM) is enabled when the rate on whichever wheel exceeds a prescribedthreshold. As the rate is reduced below a threshold, the NOM is automatically activated. In case of failure, aRate Damping Mode (RDM) based on stellar sensor rate mode is enabled (first-level fallback). Double failure(Normal mode and Rate mode) determines fallback to SAM or Safe Mercury Mode (SMM). After fallback,restart of the nominal sequence (IAS, NOM) shall be enabled from ground.High Boost Mode (HBM) is enabled from NOM, after cruise separation, to insert the spacecraft into Mercuryorbit and change the orbit plane. HBM requires integrative compensation to remove the bias due to stepdisturbance at main engine switch-on and switch-off. Attitude control cycle during boost must be as low aspossible to permit fast reaction to high disturbance torque. A control rate of 10 Hz should be sufficient. At theend of the ∆v boost, the transition to NOM is activated. Orbit Correction Mode (OCM) is enabled to correct theorbit, when requested by the ground control centre. Before thruster activation, the spacecraft is rotated to theproper attitude. At the end of correction, the transition to NOM is activated.Safe Mode during cruise is managed by enabling SAM, based on RLG and SAS measurements. AroundMercury, the spacecraft is compelled to maintain the nadir pointing attitude for thermal reasons. A gyrocompassingstrategy (RLG based propagation, MHS based correction) is suitable to manage the safe mode inthis scenario.147


BepiColomboRDMOCMRATEDAMPINGMODEORBITCORRECTIONMODEWOMSAADS (1)RCSSAADS, LARCSSAADSRWL, RCSWHEELOFF-LOADINGMODEIAMNOMHBMSAADSINITIALACQUISITIONFROM STARSSAS, RLGHIGHBOOSTMODERCSSAMNORMAL SAADSOPERATIVEMODE RWLSAADS (2) , LARCSSMMSUNACQUISITION &SAFE MODE(fallback duringcruise)MHS, RLGRWLSBMSAFEMERCURYMODE (fallback during Mercury orbiting)STAND-BYMODE(1) Rate determination capability upto about 5°/s required(2) Attitude determination rate notlower than 10 Hz is required to copewith strong step disturbance torquePerformanceFigure 6.4-25: Mode transition diagram.The system design has been verified through frequency domain and time response analyses by Matrix-x, plussimulations run on a high-fidelity simulator qualified by the experience of the BEPPO-SAX satellite. Theperformance requirements and a preliminary error budget are shown in Table 6.4-6. The most demandingrequirements come from operation of the cameras in Mercury orbit (attitude knowledge equivalent to 1 pixel,corresponding to 5 arcseconds; attitude control stability of 10 µr/s equivalent to 2 arcsecond/s). We note, inparticular, that the AME requirement is marginally met but payload and mechanical error estimates are based onsimplified models and assume in-orbit calibration to remove systematic errors, further study is required forerrors and post-flight data processing.Error Index(arcsec)APE(X, Y)Jitter(1 ms)RPE(X,Y)Short-term(1 s)Long term(1 min)Payload 5 0 0.5 1 2Mechanical (alignment,thermoelastics, microvibration)Attitude(measurement & control)AME(X,Y)11 0.5 1 2 24 0 0.5 4 4Total Error 20 0.5 2 7 5Requirement 240 0.8 2 72 5Table 6.4-6: Preliminary pointing error budget. All figures are given for a 95% confidence level6.5 MERCURY MAGNETOSPHERIC ORBITER6.5.1 ConfigurationThe configuration of the Magnetospheric Orbiter (MMO) (Figure 6.5-1) is driven by the thermal design bothw.r.t. shape and size. MMO is a spinning satellite with the spin axis perpendicular to Mercury’s equator. Theexternal shape is a flat cylinder; the side wall is exposed to the Sun and carries the solar cells, whereas the topand bottom sides are used as radiators to achieve the required thermal environment for the scientific instrumentsand the satellite subsystems.148


Spacecraft CompositeDEPLOYABLE RIGIDBOOM FOR RPWHDEPLOYABLE HIGH GAINANTENNADEPLOYABLE RIGIDBOOM FORMAGNETOMETERSDEPLOYED WIREBOOM ANTENNAStructural DesignFigure 6.5-1: Magnetospheric Orbiter configuration.The structural design (Figures 6.5-2, 6.5-3 & 6.5-4) follows an established conventional approach. The mainstructural load bearing element consists of a truncated cone manufactured from CFRP. This supports the rest ofthe structure and transmits all of the loads to the MMO satellite interface with the interim structure attached tothe cruise module. The cold gas tank is positioned inside of this cone with the centre of the tank situated aboutthe currently calculated centre of gravity. The upper mounting flange of the cone interfaces with the antennadespin mechanism. Both the equipment mounting panels are attached with flanges to the inner cone. Thesepanels are further stiffened by the addition of four shear walls that interface with the cone and both equipmentpanels and the outer half rings. The 2 equipment panels have the payload and satellite bus equipment mountedon their internal faces so that the external faces can be utilised as radiators. The two CFRP half rings areattached to the equipment panels and shear walls and act as a thermal path between the panels. These two halfrings protrude beyond the plane of the equipment panels and are used to mount the support the thermal controlsystem, both active and passive. Finally the outer ring plus solar array and thermal enclosure equipment isattached by a sub structure to the half rings. The outer ring is not envisaged as being a load carrying element butwould only support the solar cells and thermal shielding equipment. All of the scientific instruments would befixed on one of the outer half rings or directly to the equipment panel.Thermal SubsystemThe main driver for the design of the MMO is the extreme thermal environment encountered within theneighbourhood of Mercury. To maintain the temperature of the spacecraft within an acceptable range, it isnecessary to expose the minimum surface area of the satellite towards the sun whilst allowing large areas toradiate to deep space. Because of the large radiative area needed to maintain an acceptable temperature at themaximum temperature excursion, it is envisaged that a form of active temperature control will be needed tocontrol the minimum end of the temperature range. This is currently defined as two sets of louvers situated onthe +Z and –Z faces. These would be constructed of a bimetallic system that automatically opens and closesdepending on the satellite internal temperature. When the louvers would be fully open the equipment radiatorpanels would then be exposed to space. When closed the internal temperature would be maintained. Theselouvers are supported by the central cone at the inner edge and by the two half rings at the outer edge. The outerring is covered by a mixture of solar cells, second surface mirrors and thermal blanket. It is extended far enoughto keep the opened louvers in shadow at all normal satellite inclinations. The inner surface of the outer ring /149


BepiColombosolar array is exposed to allow it to radiate. Inside of this ring the main satellite structure is enclosed by athermal blanket positioned at ≈45° to reflect radiation away from the body.CommunicationsFigure 6.5-2: Main dimensions of spacecraft.For communication with Earth, a de-spun high gain offset antenna operating at X-band and two X-bandantennas covering a hemisphere are employed. Communication with the Mercury lander will be achieved by amicrostrip UHF patch antenna.It is a requirement that the high gain antenna be able to be pointed at the earth during the actual Mercurymission phase but it is also intended to use it during the cruise phase of the mission as the main link antenna tothe Earth. This means that this antenna will require a clear field of view to enable the antenna to be pointedtowards the Earth at all phases of the mission. During the cruise phase the satellite operates in a 3 axis stabilisedmode whilst during the period that the MMO is orbiting Mercury the satellite is operating in a spin stabilisedmode. This combined with the orbital inclination of Mercury generates a need for an antenna that can be pointedanywhere within a 360° rotation about the spacecraft Z-axis plus a ±12° field of view within the XY-plane. Toachieve an unobstructed field of view when in the -12° extreme of the pointing range the antenna assembly mustbe positioned far enough away from the satellite body so that the boresight of the antenna clears the upper edgeof the structure. Furthermore this antenna must be positioned about the spacecraft spin axis to enable the wholeassembly to be despun when the satellite is operating in spin stabilised mode. Finally to accommodate theantenna within the launcher envelope and to offload the spin and pointing mechanisms the antenna needs to bestowed during the launch phase and then deployed after launch.During the cruise phase the X-band antennas must be positioned so that they are not obstructed by the sunshield mounted on the cruise / lander / MMO assembly and they must be positioned at 120° to each other toachieve the maximum coverage possible.The microstrip patch antenna for communication with the lander element is situated on the -Z face of thesatellite inside of the interface ring.Equipment and Instrument AccommodationThe various scientific instruments have differing fields of view or positional requirements (Figure 6.5-5). Inmost cases it is not possible to have a totally unobstructed field of view but any obstructions have been kept to a150


Spacecraft Compositeminimum. All spacecraft control equipment such as thrusters have been positioned as far away from any opticsas possible to avoid any plume effects. The satellite bus equipment and the scientific instrument electronics havebeen positioned on the inside of the +Z and –Z panels respectively. This allows for a simpler integration and testprocedure and keeps the main communication equipment closer to the high gain antenna thus allowing for ashorter path for waveguide and/or harness. The deployable Cluster-type booms plus their associated instrumentsmust be stowed for the launch and cruise phase of the mission and have been located at the –Z face on speciallyhardened supports and holddowns. They are arranged either side of the interface cone and positioned that theycan deploy radially outwards below the lower edge of the outer ring.ANTENNA DESPINMECHANISMUPPER LOUVRES (SHOWNCLOSED) FOR THERMALCONTROLUPPER RADIATOR &SPACECRAFT EQUIPMENTMOUNTING PANELINNER SHEARWALLSLOWER RADIATOR &SENSOR EQUIPMENTMOUNTING PANELLOWER LOUVRES(ALSO SHOWN CLOSED)FOR THERMALCONTROLMAIN LOAD CARRYINGINTERFACE CONEFigure 6.5-3: Internal layout.Figure 6.5-4: Satellite exploded view151


BepiColomboA cooler and baffle have been situated on the inside of the interface cone with the aperture facing along the–Z axis. This is thermally attached to the end optics of the spinning camera via a metallic strip.Table 6.5-1 shows the Magnetospheric Orbiter mass budget.DEPLOYED SATELLITE AS SEEN FROM–Z SHOWING ALL SENSOR FIELDS OFVIEWFigure 6.5-5: Sensor fields of view.SubsystemMass [kg]Structure 32.0Thermal 15.5Attitude incl. RCS and propellant 12.0Avionics 5.5RF X band equipment 20.6UHF equipment 3.5Energy 15.8Harness 18.5Rigid boom assembly for MAG 5.0Rigid boom assembly for RPWH 2.0Positive ion emitter 2.7Total payload 24.1Total Mass 157.2Table 6.5-1: Magnetospheric Orbiter mass budget152


Spacecraft Composite6.5.2 Thermal DesignThe main features of the MMO thermal design are shown in Figures 6.5-6 and 6.5-7.DespunPlatformHigh GainAntennaGroup(Upper)Radiator(Louversremoved)SolarPanels(30% cells)DespinMechanism(Upper) Conebehind the SolarPanelExternalBeltFigure 6.5-6: MMO external thermal configuration .Equipment mountedon the upper radiatorEquipment mountedon the internal beltEquipment mountedon the external beltLoad ConeEquipment mountedon the lower radiatorFigure 6.5-7: MMO internal thermal configuration.They include (Briccarello 2000):• a cylindrical chamber, insulated on the side (normal to the Sun) with high temperature MLI, and OpticalSolar Reflectors on top of it; the chamber hosts service equipment and internally mounted payloadelements;• radiators on top and bottom surfaces (free from the Sun, exception made for the reflections from the highgain antenna equipment), covering all available area except that taken up by the despin mechanism (top)and the separation interface and the UHF patch antenna (bottom); the internal sides of the radiators arethermally coupled via high-conductivity links, and the external sides, covered with Optical Solar Reflectors,are equipped with Helios-type louvers;• body-mounted payload elements (CPD, EPD, IMS, EEA, SCAM) hosted in the interior of the chamber,with small detector heads only protruding to the outside, protected by metallic baffles with MLI (CPDexcepted); external metallic wires for the RPWE, dispensed by internally mounted wire deploymentmechanisms; the internal boxes are also linked via metallic straps to the radiators;• body-hosted and radiators-mounted payload and satellite service equipment elements; these boxes arethermally connected directly to the internal faces of the two radiators (generally, payload service elementsto bottom radiator, satellite equipment to top radiator), via bolts and thermally conductive fillers;• boom-mounted payload items (RPWH, MAG) with local coupled radiators (placed on top and bottom areasof the payload boxes); local internal detectors conductively linked to central internal electronics viabooms/thin wires (not shown in the pictures);• a central tank filled with cold gas (Nitrogen), placed inside the internal load cone, and directly linked byconduction only to the load cone itself, whose thermal conditioning is however eased by the adoption of153


BepiColombothermal properties analogous to the radiators ones on the load cone lower closure disc (apart from the patchantenna area);• 2 belts with solar panels, 30% solar cells, with backsides facing conical surfaces covered by IR MLI to easeinfrared emission reflection away to the satellite body.The objective of the design was to reach ‘normal’ temperature ranges (not exceeding 50 C) for internallymounted equipment (both service and payload items) along all the planned mission time span. Also, the thermalperformances of the structural and thermal hardware items were to be monitored, as well as those of the solarpanels, because of the extreme environmental conditions and the large temperature excursion betweenperihelion and aphelion orbital phases.The analysis cases selected to validate the thermal design were ‘Midnight periherm’ (Mercury perihelion) and‘Noon periherm’ (Mercury aphelion), which represent adequately the environment extremes, respectively as“hot case” and “cold case”. In the hot case, the solar flux amounts to 14490 W/m 2 , and its effect is mitigatedduring the ~24 min eclipse duration close to the anti-subsolar point, while in the cold case the solar flux is lower(6290 W/m 2 ) and the longer (~152 min) eclipse phase occurs far away from the planet surface.Although the objective of ‘normal’ temperature ranges (oscillations over the orbit time) for the payload,equipment and structures were achieved without variable-surface radiators, some lower limit temperatures weretoo cold. Therefore, louvered radiators were adopted, modelled after the Helios design (Beckmann 1983). Theselected louvers, which are radial blades rotating around their longitudinal axis (parallel to radiators radii), coverboth the top and bottom radiators circular coronae (not the lower closure disc), and provide an effectiveemissivity variation from 0.8 (blades fully open, radiators free to radiate to space in hot case) down to 0.1(blades fully closed, radiators prevented from radiating to space in cold case).The analysis was performed in two phases, covering separately the external elements (in a simplifiedconfiguration without louvered radiators) and the main MMO body with no external detector heads. Theanalysis included the effect of the high gain antenna on top of the upper (“north”) radiator, reflecting Sun andplanet radiation into it. The following main features characterised the two extreme cases:Cold case, at beginning of life (at planet) conditions, louvers closed:• MLI absorptivity = 0.20• OSR absorptivity = 0.10• HGA pointing at the Earth at 90º from the Sun (minimum reflection inside the radiator)• Nitrogen tank full.Hot case, at end of life (at planet) conditions, louvers open:• MLI absorptivity = 0.30• OSR absorptivity = 0.20• HGA pointing at the Earth at 5º from the Sun (maximum reflection inside the radiator)• Nitrogen tank empty.The temperature extremes found in the analysis are shown in Table 6.5-2.Solar panels, modelled including Optical Solar Reflectors, in order to provide the correct average power perorbit (eclipses included), yield: ~470 W at perihelion, ~205 W at aphelion, which matches quite well the powerdemand (~185 W ). Excess power can be partly dedicated to possible heaters, if necessary.154ItemMidnight Periherm(perihelion)Temperature range (°C)Noon Periherm(aphelion)External Belt MLI -45 120 -130 165Conical honeycomb MLI 20 215 -50 120Central belt and conical honeycomb structures 30 50 20 35High Gain Antenna Group -60 250 -135 140Radiators 20 55 25 40Low Gain Antenna -65 5 -70 15Nitrogen Tank 15 30 5 10Despin Mechanism MLI -5 125 -85 120Despin Mechanism 50 65 10 30Service equipment on top radiator 25 50 25 40Internal payload boxes on bottom radiator 20 45 25 40


Spacecraft CompositeInternal belt mounted payload boxes 25 50 25 40Externally mounted detector heads (*) 12 40 -15 20Booms (internal) (*) 0 40 -30 0RPWH, boom mounted (internal) (*) -50 -30 -70 -20MAG, boom mounted (internal) (*) -35 -20 -50 -10Solar panel (upper belt) -60 140 -90 175Solar panel (lower belt) -75 120 -90 175Table 6.5-2: MMO temperature results. Items marked with (*) are results from the intermediateanalysis, without louvered radiators6.5.3 Avionics and Electrical DesignFunctional / electrical conceptThe MMO functional / electrical concept is depicted in Figure 6.5-8. It lists all major electrical functions andconfiguration items required by the cruise and science phases of MMO.The design is based on the following concepts:• RF Communication system in X-band with a despun high gain antenna;• Central Payload Interface Unit (CPIU) for the Payload, based on experience with ROSETTA (the CPIUcollects all instrument data and distributes the secondary power from a central power supply);• separation between system and P/L tasks (the P/L shall not use the system processor as indicated bycrossing out these tasks from the conceptual diagram).Figure 6.5-9 shows the MMO Conceptual Electrical Design which defines the functions of all units andsubsystems and forms the basis for the mass and power budgets (MMO H/W tree).AvionicsThe initial mass budget revealed the necessity of miniaturising the Control and Data Management System(CDMS). For MMO, it is proposed to use a Highly Integrated Control and Data System (HICDS) within aminiaturised CDMS. The HICDS was designed by DSS as a GSTP 3 Project (Figure 6.5-10).The HICDS module will have the double eurocard format and will apply miniaturisation technologies(ASICS), horizontal and vertical multi chip modules (MCM-H and MCM-V). The design of existing moduleswill be applied to these new technologies, because they are described in a specific hardware language as 'VHDLmodels'. Intellectual Properties (IPs) for the applied ASIC VHDL Models covers:Packet TC (ESTEC), VCA/VCM for Packet TM, Virtual Channel Assembler (VCA) and Multiplexer (VCM)(ESTEC), compact PCI bus (ESTEC), IEEE 1355 interconnect links (DSS), CAN Bus Controller (licence fromBosch via ESTEC free of charge).This miniaturisation leads to estimated mass: < 0.9 kg, dimensions: 190 x 250 x 30 mm, power: 5 to 6 W.The HICDS also includes a Compact Computer Core (CCC) developed at DSS as MCM-H, (EM verified on8 February 2000). The CCC and the MMS-Velizy 'System On-A-Chip' are ESTEC- initiatives which areregarded as possible technologies for future small missions (Hybrids, MCMs, Interconnections andMicropackaging Technologies - Development and Qualification Activities - Five Year Strategic Plan, ESTEC970 720 QCT-AB, Issue 6, 16 February 1999).The HICDS will consist of 2 µRTUs (2 x 350g) and 2 AOCS boards (2 x 600g) and will form a completeminiaturised CDMS.The CDMS will be designed in rad hard technology (ASICS) and protected against SEU and latch-up (ERC32, memory error detection and correction - EDAC - with memory scrubbing and peripheral interfaces,provisions for mitigating SEU effects on functionality, e.g. filtering by hardware and software). The Mercuryradiation environment must be assessed with respect to performance and reliability.155


BepiColomboTCHK-TM& Science Data TM LinksTM from MSETC to MSEM erc ury M ag ne tospheric Orbi terRF CO M S X -BandTM/TC Equ ipmen t• Transponder X/ X• SSPA• RFDU• 2 LG As on MM O• 2 LG As on CPS (TBC)• 1 HGA XHGADespin& TiltMech.MSE C OMS UHF-Band• Transceiver• RFDU• An tenn as• MSE TM Decoder• MSE TC Encoder• MSE link acquisition• MSE trackingAO C S-Avioni csAO CS Actu ators AOCS Sensors• none, except propulsion • Su n Sen sors• Star Sen sor• Mercury Sensor• GyrosAO CS I/F• AOCS-Sen sor/Act uat. I /F.Pro pul si on SystemCold gas sys tem• TBD th rusters• Pressu re tran sducerControl ElectronicsDeplo yment &Mecha nism s• Sepa r. from CPM r• H G A releas e m ech.• H G A des pin m ech.• H G A tilt m ech .• P/L rigid booms• P/L w ire boom sTM PCMTC PCMTM/TCPyrosPowerTBDUmbilical for stackedM MO , M SE, and CPSCP S Configuration:2 active Packet TC Decoders with CPDU• High Priortity CMDs (HPCs) as Pu lse CMDsR econ figuration Modu le & Saf eguard Memory(FDIR Implementation )Avion ics & P/L Processor M odule (cold red)H K-TM & S cien c e T M Fo rm at te r• (C C SD S Pack et s & C AD U s , C o n ca t en a te d E nc o din g• D a ta R at e S e le c t io n (L G A o r H G A p e rf o rm a n ce )I/F Controllers for• S ystemTM/TC I/F (serial bu s, links, Pu lseCM Ds, An alog & Digital Status) ;• P /L Synch I /F; P/L Sen sors & Actu ators• AOCS & Propulsion I/F;.C ontrol & D ata M anagem ent Process ing Sy stemDataSto rageSolid StateMas sM emoryforScience T M& H K -TMB us/System C ontrolTasks:• Auton omy for separation• Cruise n avigation control• Attit ude and spin rate cont rol• HG A despin con trol• HG A tilt an gle control• Lan der TM/TC, ranging• Lan der data relay• Predefin ed Tim elin eexecution• MM O TM/TC & OperationsP /L ControlTas ks:• P/L M ode cont rol• P/L S ensor &Act uator cont rolon Application LayerP/L DMS Tasks:• D ata Acquisition• D ata Filtering• D ata Com press ion• D ata Form atting• M emory M anagementT/CPowerT/C S/ST M /T C• Compen s. Heaters for cru ise• Lou vers• Thermsitors• control by C&DH S/ W• Thermostat s (TBC)• Esth er (TBC)• rotat able shield (TBC)Constrain ts to U sers:• H igh tem peratu reelectronics for TBDM MO Po werS ou rc e• Solar A rrayfor SpinnerC ruis e P ow erS ourc e• Solar Arrayof S EPMPow er S/ST M /T C I/ FPow er C ontrol• Power Regulation Power• Batt ery Charge/ Dis trib.Disch arge Control & Py roBatter ySwitchedPowerto all S/SsT/C PowerPyro CircuitsI nstrumen tprimaryPowerM M O F un cti onal/Elec tri ca l C on ceptSe n s or ra wD at a• TBDS e n sor & A ctu a to rh i-le v el c o nt r o lPayload Sensors :• T BDS y nc h I/ FPa yloadP uls e C M D s , A na lo g& D igita l S tatusPayload A ct uat ors:A ll Inst rum e ntsprim ary P ow e rPowerSupplyCPIU C entral Payload Int erface U nit (bas ed on RO SE T TA PIU)S e c ond ar y Pow e rS pecific Tasks:• R igid & Wire boomrelease/deploym entFigure 6.5-8: MMO functional / electrical concept.P /L Co nfigurat ion:Magnetometer (MA G)• Dual 3-axis flu xgatePlasma Ion Spectrometer (IMS)• Mass spectrometerElectron Electrostatic Analyser (EEA ):• E lectrostatic analyserE-Wave analyser:• Dipole with anten naH -Wave analyser:• Search coils, 3 axes fieldCold Plasma Detector (CPD):• M ass SpectrometerEnergetic Particle Detector (EPD):• I on telescope, e-telescopeScann in g CameraM M O _e l_c o nc ep t,3 G . Kah l, 25 .0 2.200 0TMTCTCTMLGAsHGAH GAT iltingTM from MSETC to MS ERF COMSX-BandHICDSTRSPERC 32Antenna Dish14 to 20 MIPS• DMS tasks 0.5 MIPS• AOCS tasks 0.5 MIPSDishDespinHighly Integrated Control & Data SystemSSMMAIMAOCS SensorsAOCSAOCS InterfaceModulesCold GasPropulsionMSE COMSUHF-BandRF ComS &TrackingequipmentMSE TC Encoding& TM Decoding .equipment(S/W runs in CDMU)T /C S/SPow er(from PD UH eaters /L ouv ersT herm is to rsI/O Channels2 µRTUs(embedded in CDMS)TM /TCDeployment &Mechanisms• Separ. from Lander• HGA release mech.• P/L rigid booms• P/L wire boomsP yrosPow erTBDData Bus e.g. CAN BusAll Instrumentsprimary PowerSolarA rrayPCMPCM PCDULi-Ion BatteryPDMSwitched PowerPermanent PowerSwitched T/C PowerPyro CircuitsC CSDS DataPacke tsData Bus RTTM /TCSynch I/FCPIU Central Payload Interface UnitP u lse C M D s, An a lo g& D i git a l S t at us(based on ROSETTA PIU)PowerSupplySEPMSolarArra y(cruise)Power S/SE-Wave AnalyserMagnetometerPlasma Ion SpectrometerSecondary PowerScanning CameraCold Plasma DetectorH-Wave AnalyserElectron ElectrostaticAnalyserEnergetic Particle DetectorMMO Conceptual Electrical DesignFigure 6.5-9: MMO conceptual electrical design.156


Spacecraft CompositeCAN BusGPIOCLCKCCCSystem SerialNetworkIEEE 1335Compact PCI BusDRAM controllerDRAMe.g. 640 MbitChannelMultiplexerand TMencoderPacketTC DecoderHPCDriver64High PowerCommandDataChannelsDownlinkUplinkFigure 6.5-10: Highly Integrated Control and Data System (HICDS).The mass and power of the miniaturised CDMS are given in Table 6.5-3 and compared with a state-of-the-artCDMS (LISA is taken as reference).State of the Art CDMU (LISA)embedded modulesMiniaturised CDMU2 HICDS, 2 AOCS I/F boards, (2µRTUs virtually inbudgets ), 2 DC/DCMass / kg 15.9 2x0.9 + 2x0.6 + 2x0.35 +0.6 + 1.2 = 5.5Power / W 25 12Dimensions /mm 3 410 x 240 x 190 1200 x 250 x 190MMO Data Return BudgetTable 6.5-3: MMO CDMS mass and power.The following budget gives the specified data rates of the MMO experiments (Table 6.5-4), the data volumeof the MSE data relay (4176 kb per orbit, with a data rate of 8.7 kb/s and a contact time with MSE of 480 s forthe first 7 days of the MSE mission). The S/C HK data rate has been limited to 200 b/s.To comply with the constraint of the mean X-band link performance of 211 Mb, the specified mean data ratesof the instruments must be weighed with a certain duty cycle (dc). The mean link performance is specified for33% contact time of the orbit to the GS at a rate of 18.2 kb/s. The real telemetry rate varies between 9 kb/s(maximum distance to Earth) and 65 kb/s (minimum distance).The duty cycle has not yet been finally defined. The given budget is an example, it could be kept at 100% forall instruments duty cycles except for e.g. the spinning camera. The optimum dc of the camera in this case couldbe 62% which would still correlate with a camera data volume of 107 Mb per orbit.The calculated budget figures thus result in a mean data return of 210.8 Mb per orbit. This is the net figurefor the user data, overhead for encoding and packetisation is respected in the link budget.PowerThe power figures are given in Table 6.5-5. With the design efficiencies of a Maximum Power point Trackerand Main Regulator, this profile requires a total energy of 132 Wh for a long eclipse at aphelion and a solararray power of 186 W in the Sun (for 266 mn), when the transmitter is off. During data transmission, the batteryshould not be charged to avoid power dissipation in the S/C because of the efficiency losses of the solid statepower amplifier (SSPA) of the COMS and battery charge regulator. The maximum dissipation for radiatordesign thus can be limited to 131 W during data transmission in sunlight.157


BepiColomboS/S Unit Name Abbr. Data R. Data R. Data R. Duty Data R.AvionicsUHF EquipmentPeak Average Min. Cycle Aver. calcul./bps /bps /bps (DC) /bpscommand & data mgt unit CDMU 200 1 200TM from Lander (4176kbits per orbit) 8700 1/72 121Positive Ion EmitterMagnetometerPIEelectronics 20 20 20 1 20MAGelectronics 1200 800 100 1 800Plasma Ion SpectrometerIMSIMS unit tbs 500 tbs 1 500Electron Electrostatic AnalyserEEAEEA unit tbs 100 tbs 1 100Wave Analyser (E)Wave Analyser (H)RPWERPWHWave Analyser Common Electronics 10000 1000 100 1 1000Cold Plasma AnalyserCPDCPD unit 400 200 20 1 200Energetic Particle DetectorEPDEPD unit 100 40 20 1 40Camera for Spinning S/CSCAMelectronics tbs 5000 tbs 62% 3100Central Payload Interface UnitCPIUCPIU 100 20 20 1 20Sum of Data Rates /bps 6101Orbit Period /s 34560Data Volume per Orbit /Mb 210,8Dump Capability per Orbit /Mb * 33,60% 18200 211,3Margin per Orbit/Mb 0,5*overhead: 15% RS, 2% packetisation are respected in link budgetTable 6.5-4: MMO Data BudgetPower Profile Mode inAphelion:Eclipse Sun,Tx on Sun, Tx off,Batt. chargemean dissipation/W 119.7max dissipation/W 131.55power demand/W 55.4 158.36 139.36time/min 118 192 266,4Energy required132WhSolar Array Powerrequired171W186WTable 6.5-5: Worst case power profile (at aphelion)Scaling down the Li-Ion battery of Mars Express to 3.5 kg provides 186 to 260 Wh of stored energy(corresponding to 70% up to 90% battery efficiency, end of life, depending on source resistance, load,temperature, and charge status of battery). Taking into account a one failure tolerance of the 12 strings with 6cells in line each, this battery could provide the required 132 Wh at 80% DOD.The PCDU is designed with the mass and dimensions given in Table 6.5-6.158


Spacecraft CompositePCDU Total Mass12.0 kg10.3 k gDimensions:L x W x H (mm x mm x mm)475 x 204 x 203375 x 204 x 203Table 6.5-6: PCDU, standard design and mass optimisedThe Solar Array of MMO must be designed for aphelion with temperatures of 105 C for the upper part and 85C for the lower part (effect of the large antenna). For a power demand of 186 W the SA projected area withstandard GaAs cells must be ca. 0.233 m² (upper area), corresponding to a cylindrical area of 0.732 m 2 (∅ 1.7 mx height of ca. 13.7 cm). The corresponding figures for the lower SA are 0.223 m² projected area and 13.1 cmheight because of its lower temperature.6.5.4 CommunicationsThe telecommunication elements in the MMO spacecraft comprise:• Earth telecommunication elements (X-band), to provide tracking, telemetry and commands in the Mercuryorbit phase. In the split-launch option, the MMO Earth TLC elements are also used during the cruise; in theAriane 5 option, they are off.• Data relay telecommunication terminal (UHF-band) for communication with MSE (split-launch optiononly).Earth telecommunicationsThe main features of the MMO Earth telecommunications scenario are the same as those described in Section6.4.4 (MPO). The differences are: (1) planet occultations are negligible for the high MMO orbit, and (2) theantenna pointing mechanism has no elevation or azimuth constraint, provided that the elevation degree offreedom can be implemented by RF means. The assumptions are the same as for MPO (Perth station, 35 m dish,available for 8 hours out of 24 with spacecraft elevation > 10°).The proposed system is X-band up/down, since only limited radio science experiments are planned withMMO and the data collection rate is smaller than for MPO. The proposed on board system architecture includes:• redundant X-band chain for telecommand reception and telemetry transmission (Mercury orbit + cruise inthe split-launch option)• redundant X-band transponder• 20 W RF output power HPA’s• 1 m High Gain Antenna reflector, 39.5 dB peak gain, despun and ±11° elevation excursion• 2 Low Gain Antennas (X-band) with 9.5 dB boresight gain, 0 dB at 50°.The normal HGA operation mode is residual carrier, including ranging. The available telemetry rate variesbetween 9 kb/s (maximum distance to Earth) and 65 kb/s (minimum distance). During the mission, the telemetryrates are varied in discrete steps to cope with the varying distance to the spacecraft. Figure 6.5-11 shows thetelemetry rate as function of time at X-band, assuming 20W RF power and a 1m antenna reflector, in theresidual carrier mode.Data rate [kbps]50454035302520151050804.618821.886Data rates at X Band840.488854.387895.915921.428932.887973.9791017.9561029.906Mission elapsed time [days]1046.4941078.9771091.7661142.3281165.813Figure 6.5-11: Telemetry rate as function of mission time (X-band, residual carrier).159


BepiColomboThe data returned to Earth by the telecommunications system is calculated from the above figures, aftersubtracting the time when the ground station link is not available. The estimated data return after 1 year is 160Gb.High Gain Antenna configuration and design. The design drivers of the High Gain Antenna subsystem are theenvironment (high operative temperature and temperature excursions; dimensional stability; pointing stability;radiation) and the RF design (X-band, radiating aperture limited to 1 m, constrained by spacecraftaccommodation). A possible configuration of this antenna is shown in Figures 6.5-12 and 6.5-13 (ESTEC,1994).The requirement for the steerability of the beam is fulfilled with a tiltable subreflector. The innermostreflector is tilted via a small actuator mounted on the despun platform. The antenna configuration is derivedfrom a dual-offset Cassegrainian configuration. There are three reflectors in the antenna design. The mainreflector is a fixed paraboloid; the second subreflector is a hyperboloid and is also fixed. The design is not anopen Cassegrainian system and, as a direct consequence of the positions of the reflectors, some apertureblockage due to the first subreflector is unavoidable. This option has been chosen to preserve designcompactness, and therefore losses have been taken into account. The total angle to be scanned is 24°, thusimplying the need for 10 different selectable positions for the tiltable subreflector. Choosing a 1 m mainreflector and 2.50° beamwidth we obtain 35.88 dB as a first result for the gain of the antenna (efficiency is set to0.5). This includes losses due to the diffraction effects of the satellite structure, spillover and cross polarisation.Figure 6.5-12: MMO antenna layout.160Figure 6.5-13: Antenna reflectors.A detailed study is required in the subsequent phases with highly specialised tools, such as GRASP8 package.A despin mechanism is required to point the antenna towards Earth. The configuration has the advantage ofcoupling the satellite with the spinning platform without any device to transfer the RF power to the feed. Theonly exception is the transfer of electrical power to the little actuator on the platform.The RF waveguide is located along the spin axis of the satellite and illuminates the tiltable subreflector with acircularly polarised wave. The circular polarisation of the RF signal is mandatory, because this avoids that thespin of the platform interferes with the power radiated from the feed.


Spacecraft CompositeThe thermal performance of the antenna is optimised by using a grid structure for the main reflector. The dishsurface is realised by stretching an RF reflective metallic mesh, e.g. a thin gold-plated molybdenum grid(partially transparent to the solar flux), onto the antenna structure. Using a metallic mesh instead of a solidsurface, it is possible to reduce the effective reflecting area to one tenth of the surface, maintaining the same RFperformance. Materials for the supporting structure shall be similar to those adopted for the X/Ka band antenna.Data relay telecommunicationsA relay between the Magnetospheric Orbiter and the Surface Element is implemented in the split-launchscenario. Data are collected by the Surface Element at low rate and transmitted to the Orbiter at each overheadpass. The generic design requirements and assumptions were the same as in the MPO case (Section 6.4.4), withthe exception that Turbo coding of the MSE data increases the link capacity. The design assumptions and linkanalysis results are given in Table 6.5-7.Data rate / LifetimeLander output RF powerAntenna gains @ θ = 40°Limit slant rangeDuration of comms windowSurface Element to MMO8.7kbps / 7 days1WNumber of passes in 7 days 18Data return in 7 daysOrbiter 2 dB, Lander -1 dB2000 km8 min75 MbitTable 6.5-7: Data relay assumptions and link analysis resultsFor the Lander to MMO link, the useful window is identified at slant range ≤ 2000 km, when each elementhas the other within 45° of its antenna pattern symmetry axis. The contact time lasts 8 minutes per orbit (Figure6.5-14), and the data rate available from the link budgets is 8.7 kb/s with 1 W MSE RF power (includingcoding). With respect to the MPO case, the communications window has the same length (8 min), but the link isless efficient because of the small number of passes, due to the longer orbit period, and the longer range (2000km assumed as upper limit). This is only partially compensated by doubling the RF power. Overall, in the MMOoption the data budget collected over 1 week is about half the budget in the MPO option.9080Slant range (100km) & elevation (deg)706050403020100530 535 540 545 550 555 560Time from pericentre (min)Figure 6.5-14: MMO – Lander communications window. The continuous line is the slant range, inunits of 100km. The dashed line is the elevation above the Lander horizon, in degrees. Dotted verticallines identify the permitted communication intervals.For the Magnetospheric Orbiter, configuration constraints lead to selecting a flat antenna. A microstrip patchantenna is tentatively chosen. Helical quadrifilar or volute antennas would provide better performance if theycould be accommodated.This antenna is made of a metal ground plane, a KOREX (high temperature honeycomb, especially designedfor antenna applications) dielectric substrate and four conducting patches. Its mass is 1.5 kg. and the sense ofpolarisation is selectable by choosing the right feeding sequence of the patches. Its flat structure and161


BepiColomboperformance meets the mounting constraints on the lower surface of the Magnetospheric Orbiter (Figure 6.5-15).25 cm.25 cm.3.5 cm.Figure 6.5-15: MMO UHF antenna.The data relay assembly consists of a directive UHF antenna and a UHF Transceiver. The transceiverincludes Receiver/ Demodulator / Transmitter, integrating in a single box all RX and TX functions and theinterface with the Data Handling subsystem. The transceiver dimensions are (mm) 140 x 106 x 32 (off-theshelf),the weight is 1 kg. Two units are accommodated for redundancy.6.5.5 Attitude ControlIn Mercury orbit, the nominal attitude of the MMO is spinning around an axis perpendicular to Mercury’sequator, at a rate between 4 and 15 rpm (reference for the design: 15 rpm). The moments of inertia are shown inTable 6.5-8.Principal moment of inertia Booms stowed [kg·m 2 ] Booms deployed [kg·m 2 ]I ZZ 49 260I XX 21 30I YY 24 170Table 6.5-8: MMO moments of inertia. The Z-axis is the nominal spin axis; the X-axis is along thedirection of the wire booms (30m length assumed)Once the nominal attitude achieved, there is no requirement for any attitude manoeuvres other than thoserequired to counteract perturbing torques. The yearly elevation motion of the high gain antenna (±11° withrespect to the equator) is insured by the antenna itself, although an angular momentum budget for suchmanoeuvres to be performed by the AOCS has been calculated (see below). Thus, the requirements of theAOCS are limited to:• acquisition of the nominal attitude after (spinning) release from the carrier and spin-up after boomdeployment;• attitude maintenance under perturbing torques (gravity gradient, solar and planet radiation pressures).The maximum perturbing torque is due to the gravity gradient (2 x 10 -4 Nm at pericentre); however thegravity gradient is symmetric and does not contribute to the angular momentum. The solar radiation pressurewould contribute a secular angular momentum build-up is the satellite were not spinning; the effect of the spin isto null the angular momentum build-up component, leaving only a nutation angle. The maximum amplitude ofthe nutation angle, driven by the solar pressure torque, is calculated as ∼2 x 10 -4 degrees, hence negligible. Theamplitude of the nutation angle also imposes requirements on the off-diagonal terms of the inertia matrix. Forthe nutation angle not exceed 0.1º, the off-diagonal terms must be: Ixz < 0.3 kgm 2 , Iyz < 0.1 kgm 2 (designrequirement).The re-orientation of the spin axis can be executed by a set of 2 thrusters (+2 for redundancy), placed atpositions 180° apart along the outer rim of the cylinder bottom surface, with thrust vectors parallel to the spinaxis. Two thrusters are used to augment the control effectiveness in the early mission phase after release fromthe carrier, should it be needed (spin-eject release at low rate). The maximum demand for spin axis reorientationwould occur if the satellite were required to follow the yearly declination variation of the Earth, inorder to point the high gain antenna. With actuation by 100 ms pulses of a 0.2 N thruster, the required propellant(Nitrogen) demand is 0.5 kg. Although this is not the reference solution for antenna pointing, this small amountof propellant is kept as an allowance for any spin axis re-orientation manoeuvres (e.g., to correct deploymenterrors).For the spin-up, we make the worst case assumption that all of the nominal spin rate (15 rpm) must beimparted by thrusters. The redundant layout for this task includes two pairs of two thrusters, placed at positions162


Spacecraft Composite180° apart along the outer rim of the cylinder, with thrust vectors tangential to the cylinder. With F = 0.2 N,1200 s are required at propellant consumption of 0.8 kg. No propellant at all would be required if the satellitecould be spun up to 80 rpm at release; the deployment of the booms would then bring the rate down to thenominal value (ratio of deployed to stowed spin axis moment of inertia ∼5).Summarising, the actuation is provided by a set of cold-gas (Nitrogen) thrusters as the simplest solution. Aredundant thruster layout for the attitude and spin up requirements includes 8 thrusters in two blocks of foureach (shown in Figure 6.5.5). The propellant required is 2 kg for 1 year, with 50% margin. Given the lowactuation requirements, alternatives to the thruster based attitude control should be considered in the future as afurther mass-saving measure.The sensors set includes two (redundant) slit sun sensors and one (redundant) Mercury horizon sensor. Thesensor environmental requirements would be the same as those for the corresponding equipment on MPO, and acommon technology research programme is envisaged.6.6 MERCURY SURFACE ELEMENT6.6.1 Lander Based on Solid PropulsionThe hard lander variant of the Mercury Surface Element (MSE) currently has two versions (Figure 6.6-1,Table 6.6-1):• equipped with two motors (insertion and descent) for application in the Ariane 5 scenario;• equipped with descent motor only, for application in the split launch scenario, where the insertionmanoeuvre is performed by the Chemical Propulsion Module.Two-motor versionFigure 6.6-1: MSE Assembly.Single-motor versionFigure 6.6-2: Aft body on the surface with Fore body buried in the ground.163


BepiColomboWith the exception of the insertion motor and its interface structure, the basic design of the MSE is identical.Hard Lander design conceptThe Mercury Surface Element has an Aft and a Fore body as its central core. The Aft body contains all thegeochemical experimentation, cameras, power supplies and communication. The Fore body holds thegeophysical experiments and is connected to the Aft body via an umbilical chord. One (split launch option) ortwo (Ariane launch option) solid propellant motors are attached to the Aft body structure. The smaller is used toinsert the MSE into a ballistic entry trajectory, the second controls the descent velocity of MSE. At a givenheight, lateral thrust motors rotate MSE so that it is orientated vertically, and then MSE is allowed to free fall.The axial crushing of the main motor casing controls the Aft body deceleration load, whilst the Fore body is freeto impact and penetrate the surface of Mercury under its own inertia (Figure 6.6-2). The overall dimensions areshown in Figure 6.6-3.In the split-launch scenario, the intended spin of 2 Hz (synchronisation with the acquisition rate of the radaraltimeter) will be imparted into CPM prior to separation of MSE as an aid to trajectory stabilisation. CPM willalso perform the insertion burn. In the Ariane scenario, the spin-up will be imparted by a dedicated spin-ejectdevice and MSE shall, at a pre-determined point, be inserted into a ballistic approach trajectory using the smallinsertion motor.Insertion motor. The MSE insertion motor consists of a solid propellant rocket attached to the Aft bodyelement via a spin eject device. The insertion motor uses a carbon fibre motor casing and nozzle whosecharacteristics are presented in Table 6.6-2.(a) two-motor version (Fore body not shown)(b) single-motor versionFigure 6.6-3: MSE overall dimensions.Once depleted, the insertion motor shall be ejected from MSE using a mechanical delatching / spring ejectiondevice located at the interface between the motor and top of the MSE Aft body. This design is in existence, andhas been developed for similar space applications by HEL (Figure 6.6-4). The eject mechanism has a total massof 2.0 kg, and is configured so as to impart zero ‘tip-off’ into the separating systems.164


Spacecraft CompositeDescent motor. The descent motor is required to reduce the MSE approach velocity from 3.2 or 3.5 km/s to 0at a horizontal approach height of 1.5 km above the surface. The descent motor is a solid propellant rocketmotor rigidly attached to the Aft body. The characteristics for both descent scenarios are in Table 6.6-3.ItemSingle Motor VersionMass [kg]Two-motor VersionMass [kg]Insertion motor casing 0.0 4.8Descent motor casing 20.0 18.0Thrusters 11.0 11.0Control unit 3.7 3.7Radar altimeter 0.6 0.6Penetrator & support 4.2 4.2Platform 2.1 2.1Cover 1.8 1.8Support webs 0.8 0.8Total payload 6.1 6.1Antenna & comms 3.2 3.2Antenna restraint 0.1 0.1Harness 2.9 2.9Batteries 4.5 4.5CPU 1.8 1.8Thermal protection 2.1 2.1Launch restraint 3.3 3.3Motor Separation Device 0.0 2.0Insertion motor propellant 0.0 21.6Descent motor propellant 151.3 139.7Total Mass 219.5 234.3Table 6.6-1: MSE mass budget (Mass items include margins according to item maturity)Insertion ∆vSpecific impulseMean thrustBurn timeMotor diameter / lengthPropellant massMotor casing massAriane launch option0.209 km/s273 s12140 N5.0 sØ 0.27m / 0.441 m21.6 kg4.8 kgTable 6.6-2: Insertion motor characteristicsThere is the potential to switch on and off the motor using ‘hybrid’ rocket motor techniques to facilitate amore accurate control of MSE, prior to its final manoeuvre and vertical descent onto Mercury. This technologyis currently being researched by HEL.To perform this manoeuvre, MSE must possess a degree of autonomy to monitor and control its height abovethe planet as well as its velocity relative to the surface and roll rate / attitude.165


BepiColomboMonitoring the descent rate and altitude shall be performed by means of a small light weight (0.4 kg) radaraltimeter (Figure 6.6-5). The current design of altimeter is capable of sensing from a range of 4.5 km, but with aheight tolerance of approximately ±200 m. The suppliers have stated that the operational height can be extendedto 40 km, but with a mass (0.6 kg) and measurement tolerance (± 2.0 km) penalty. However, any initialtolerance shall improve with reduction in MSE approach height. The altimeter is a stand-alone unit, with astandard RS232 interface. It shall be switched on at a pre-determined time and continually generate a radarbeam (4.3 GHz) over a cone angle of 70°. It samples every 10 Hz and with an MSE spin rate of 2 Hz, and iscapable of producing a height reading every revolution of the MSE. With a calculated burn time of 26.0 s thisresults in 52 data points which is sufficient to predict the MSE approach descent rate.Figure 6.6-4: Spin up eject mechanism.Figure 6.6-5: Radar altimeter.Ariane launch optionSplit launch optionApproach velocity 3.2446km/s 3.458 km/sMean flight-path angle 8° 12°Specific impulse 290s 290sMean thrust 14765 N 14765 NBurn time 24.5 s 26.3 sMotor diameter / length Ø0.56 m / 1.07 m Ø 0.56 m / 1.168 mPropellant mass 139.7 kg 151.3 kgMotor casing mass 18.0 kg 20 kgTable 6.6-3: Descent motor characteristics.Velocity and attitude can be monitored by use of low cost / low mass solid state rate gyros. The main CPUshall be supplied with sufficient information to detect the height of the MSE and rate of change of velocity, aswell as imparting any additional stabilisation manoeuvres required to prevent out of plane misalignment with themain velocity vector. A mass budget of 5 kg has been included for additional flight control thrusters (such asdiscrete squibs). The CPU supplied with this information, referenced against an in-built descent model, willhave sufficient information to predict and detect the point at which MSE approaches the required rotationwindow.MSE is then required to rotate through ∼80° from a horizontal to vertical position. This manoeuvre can beaccomplished by the use of discrete thrusters – radially orientated around the base of the motor case. The powersupply for activating and controlling individual thrusters shall be provided by a separate module (3.7 kg).The mass prior to impact is 63.8 kg, which from a free fall drop height of 1.5 km equates to an impact energylevel of 351 kJ at an impact velocity of 105 m/s.Upon impact, the spent motor casing / nozzle will be utilised as the main energy absorption device. It shall beconstructed as a filament wound carbon composite, whose thickness (2 mm) and fibre orientation will beoptimised to withstand the internal pressurisation during motor burn and yet readily deform under the axialimpact loading. The collapse force shall be governed by the 2500 m/s² deceleration loading imposed on the Aftbody. Similar devices have been constructed and tested at HEL (Figure 6.6-6).166


Spacecraft CompositeAft body. The Aft body is constructed from aluminium honeycomb, skinned with carbon fibre, making thestructure light weight with high stiffness. All the scientific equipment, control and communication packages areattached to the main platform (Figure 6.6-7). The base of the equipment platform is connected to the descentmotor casing. The main structural support is through three radial webs, which support the external skins and theinternal core supporting the Fore body penetrator.Figure 6.6-6: Motor casing impact crush.The micro-rover restraints are incorporated into the Aft body, and their release simultaneously activatedalong with the opening of the access panel located in the sidewall and the deployment of the communicationsantennae. The micro-rover is free to exit the Aft body but remains connected to MSE by an umbilical cable(Figure 6.6-8). The camera will be positioned on the top of the Aft body cone, inside the well of the Fore bodypenetrator support structure.Fore body. The Fore Body shall be retained within the Aft body central support hub –enveloped by the mainrocket initiator. Due to high thermal loads, the penetrator shall be protected by an additional lay of intumescentmaterial which has the capacity of limiting the temperature of the penetrator to below 100 C.The penetrator shall be manufactured from a hollow titanium into which are located the geophysics packages(Figure 6.6-9). The Fore body is connected to the Aft body data handling and storage equipment by means of anarmoured spoolable umbilical chord.Figure 6.6-7: Equipment platform on Aft body.Figure 6.6-8: Deployed Aft body.167


BepiColomboFigure 6.6-9: Fore body.As the Aft body decelerates on impact, the inertia forces acting on the Fore body are more than sufficient tofail the shear-limited restraints (nominally set at 500 m/s²). Once separated, the Fore body is free to pass throughthe deforming motor casing / nozzle, then impact and penetrate the surface of the planet. Its impact orientationbeing essentially governed by its original velocity vector and angle of incidence. The penetration performanceof the Fore body into regolith is presented in Figure 6.6-10. It shows that from an original free fall height of 1.5km, the penetrator is capable of achieving depths of 7.3 m, with a mean rigid body deceleration loading of 2400m/s².Specific areas of development.The following areas within the MSE require further development:-• space qualification of altimeter and functioning;• thermal protection of Fore body;• armouring of umbilical connector;• improvement of descent motor performance (Isp);• tailoring of impact characteristics of descent motor casing;• ability to switch off descent motor (hybridisation);• characterisation of solid state rate gyroscopes.MSE Electrical SystemFigure 6.6-11 shows the major functions and configuration items of the MSE functional/electrical concept.Some of them have been deleted and are crossed out.The hard lander system uses a separate controller during the decent for the AOCS sensors and the propulsionsystem (indicated by the deletion of the related bus/system control task and I/F between the propulsion systemand the system controller).After analysis of the energy and power budgets, and under very high mass constraints, it was decided tosupply MSE with only a primary battery of high energy density. The risk of shadows on the solar array at alatitude of around 85° is high (40%); an optional small solar generator (ca. 0.5 kg) could feed directly thepayload, or part of it, without charging or discharging a secondary battery. This would keep the mass of thepower sub-system minimum.Electrical design (Power and CDMS). The initial design budgets are based on the ROSETTA landercharacteristics listed in Table 6.6-4.ROSETTA lander power from Power S/S Specification Document (RO-LPO-SP-380001):• Primary battery. 800 Wh, 2950 g incl. packing + harness (960 Wh at 1 mA discharge current, 710 to 730Wh at 20 C and 15 W discharge rate, SAFT Li-Thionyl-Chloride (Li/SOCl 2 ), derivative of the Huygensbattery, 3 cells, gel-like electrolyte, shock and 35 g rms vibration tested;• MPPT power regulation concept, 5 W to 15 W solar generator power between 3 and 2 AU;• PSS power conditioning /distribution occupies 9 boards (100 mm x 140 mm in the common electronicsbox).168


Spacecraft CompositeROSETTA lander CDMS:• redundant boards in common e-box: 1.7 kg• CDMS-Power at descent: 3.3 W; energy 3.3 Wh (1 h descent assumed)• CDMS-Power for first 56 h: 3.3/0.2 W; energy 98 Wh• CDMS-Power for next 64 h: 3.3/0.2 W; energy 112 Wh.MSE equipment proposed/assumed. Table 6.6-5 gives the estimated mass of the MSE electrical items basedon the ROSETTA lander figures. The battery mass is scaled down to 2 kg (Lithium-Thionyl-Chloride battery).A thruster control unit of 3.5 kg could replace the CDMS; the equivalent mass would be saved but power wouldbe required during the total descent phase of ca. 2h. Thus a descent battery would also be needed. The harness &balance mass is based on the ROSETTA lander experience.Mercury KEP (Rigid Body Acceleration vs Time)25002000Penetrator Performance(Crusty/Cloddy Sand)Mass Velocity Acceleration1.34 kg 60 m/s 1200 m/s²1.34 kg 150 m/s 2400 m/s²Acceleration (m/s²)150010005000Mass = 1.34 kgVelocity = 60 m/s0.00 0.02 0.04 0.06 0.08 0.10 0.12 0.14Time (s)Mass = 1.34 kgVelocity = 150 m/sMercury KEP (Penetration Depth vs Time)876Penetrator Performance(Crusty/Cloddy Sand)Mass Velocity Penetration1.34 kg 60 m/s 1.8 m1.34 kg 150 m/s 7.3 mMass = 1.34 kgVelocity = 150 m/sDepth of Penetration (m)54321Mass = 1.34 kgVelocity = 60 m/s00.00 0.02 0.04 0.06 0.08 0.10 0.12 0.14Time (s)Figure 6.6-10: Penetrator performance.ROSETTA Lander Item Net Mass/kg Margin Mass/kg Gross Mass/kgPower system / common e-box 9.35 0.45 9.80CDMS 1.70 0.00 1.70Harness, balancing mass 2.50 0.30 2.80Telecoms Units 2.80 -0.48 2.32Table 6.6-4: ROSETTA Lander mass budget for comparison.169


BepiColomboTCTMData RelayRF COMSUHF-BandTM /TC Equipm ent• Transpo nde r• S SPA• L G ATM P CMTC PCMA O C S-A v ion icsA O CS A ctu a to rs• n one , except pro pulsio nAO CS I/F• AO CS-Se nsor/Actua t. I/FAO CS S en sors• La ser A ltim eter• Sp eedo me ter• Gyros (TBC)• A cce le rom ete r (TBC)• M ercury S ensor (TBC).Propulsion SystemSolid Roc kets• insertion rockets• d ecelera tio n rocketsControl Electronicswith TimerI/F deletedDeployment &Mechanisms• Rover release• jettison of insertion rockets (tbc)TM/TCPyrosPow erTB D(Composite?) Umbilical tbcCPS Configuration:active TC De co der• enco ded co mm and s (TBC)• Pulse CMD sController or Processor Mo dule (TBC)H K-T M & Science T M Form atter• (C CSD S Packets & CADUs,Concaten ated Encoding t bc)I/Fs for• System TM/TC I/F (seria l b us, links,An alog & Digital Sta tus) ;• P/L Synch I/F; P /L Sen sors & Actuat ors• AO CS & P rop ulsion I/F;Sys tem Controlle r or µP .DataStorageSolid S tateMas sM emoryforScience T MBus/System ControlTasks:• activa tion afte r se paration• M SE command distribution• P red efined sequ ence forinsertio n ro cke tr b urn,dece le rat ion ro cke t w ithaltimete r log ic fo r cut-off• deploy antenna, SA, andaft-body payload bayP/L ControlTasks:• P/L Mode commandingaccording toreceived TCsP/L DMS Tasks:• Data Acquisition• Data Formatting• Memory Managem.T/CPow erT M/TCT/C S/S• Compens. Heaters/ Louvers• Thermsitors• Thermostats (TBC)Constraints to users• High temperatureelectronics for TBDP ri m a ry P ow e rS ourc e• only option:Solar ArraySolar Ar rayDescent Ph ase(TBC)• • •Descent Phase PowerPow er S /SPow er Control• Po we r R eg ula tio n• Ba ttery Ch arg e/Discharge Con tro lT M / T C I /FP rimary BatteryCentralPowerConve rsion,Dis tribution& P yrosPrimary BatteryactivationPrim aryPow er to S /SsPyro Circu itsSe cond aryPo we r toIn st rum en tsExperimentDataExperimentcontrolCameraMagnetometerSeismometerMicro RoverMößbauerAlpha X-RaySpectrometerAft BodySecondary PowerP u lse C M D s , A na lo gS y nch I/ F& D i gi ta l S t at usPaylo adFore BodyHeat Flow & Phys, Property Package (HPThermal SensorsAccelerometerDensitometerSensorSourceHP 3 Electronics3 )MS E Fu nc tio na l/Ele ctrica l Co nc e pt , H a rd L an derMSE_el_concept,4-hard Lander G. Kahl, 29.2.2000Figure 6.6-11: MSE functional/electrical concept.170MSE Equipment Mass RemarkThruster Control Unit 3.5 kg usable as CDMS and integrated in the Aft bodySolar Array & Support 0.5 kg option only (see note)Harness 1.0 + 1.8 kg with balancing mass (ROSETTA Lander)Batteries 2.5 kg 2.00 kg primary battery (+0.5 kg impactprotection)Power DC/DC converterPower LCLs & pyros1.0 kg0.8 kgCDMU 1.7 kg ROSETTA LanderComms 2.5 kg COMS Rx/Tx, no antenna includedTotal Mass13.3 (13.8) kgTable 6.6-5: Preliminary MSE electrical systems mass budgetNote: A complete power subsystem with SA and rechargeable battery would weigh 4.75 kg (2 kg forcharge/discharge regulation +1.25 kg secondary battery +1.5 kg MPPT & power regulators). The option withsolar cells but without secondary battery and charge/discharge regulation must be analysed in more detail).The energy budget in Table 6.6-6 is derived from the ROSETTA lander battery (710 Wh at 15 W dischargerate and ca. 3 kg). The energy available at MSE would be 473 Wh. COMS, CDMS and auxiliary equipment willconsume ca. 127 Wh, leaving 345 Wh of primary energy to the P/L.The CDMS relies on a CMOS technology and is assumed to require a power of 3.3 W for 20 h of activitywithin a 7-day life time (including 10 mn of operations prior to impact). During the remaining 148 h, the clockrate is reduced to stand-by, which leads to a consumption of 0.2 W. Alternating active and stand-by periodsmust be controlled by timers. The preliminary energy budget of the experiments is given in Table 6.6-7.MSE Data Return. Table 6.6-8 gives the worst-case data volumes (MSE to MMO link, see Section 6.5.4).The table shows the data transmitted at each orbit from MSE to MMO during a contact time of 480 s at a mean


Spacecraft Compositeusable data rate of 8.7 kb/s (column 3) and in the 7 day mission (column 4). The P/L yields 3765 kb per orbitwhich adds up to 67.968 Mb for the 7 days of operations (18 contact periods of 480 s).ItemPower(W)Total time(h)SecondaryEnergy(Wh)PrimaryEnergy (Wh)Capacity of 2 kg battery, (Li/SOCl 2 ) 473.0UHF COMS (18 x480 s/orbit = 2.4 h) 5 7.05 15CDMS active 3.3 20 66.0CDMS stand-by (CMOS low clock rate) 0.2 148 29.6Aux Power (advanced technology) 0.1 168 16.8Instrument:ConverterEfficiency75%Total Payload 259.0 345.6Margin 0.0Table 6.6-6: MSE energy budget for 7 days with a primary batteryEnergy /Wh Energy /Whsecondary primaryItem power power Referencecapacity of 2 kg battery (Li/SOCl2) 473 Rosetta Lander, Li-IonUHF COMS (18 x 480s / orbit = 2.4 h) 15 5W Rx/TxCDMS active (20 h x 3.3 W) 66 3.3W activeCDMS stand-by (CMOS), 148h x 0.2 W) 30 0.2W stand-by modeAux Power (advanced technology) 17Available Energy for Experiments: 345Experiments: specified: 75% 75% converter efficiencyMicro-Rover with experiments 50 67 Brueckner 10.3.00HP 3 50 67 7 d continous operationDecent Camera (CLAM-D) 2 3 tbcSurface Camera CLAM-S 5 7 tbcSeismometer 118 118 tbcMagnetometer 84 84 tbcMargin 0Table 6.6-7: Preliminary energy budget of the experimentsS/SDataAcquis.Time/hData Volume/kbitsper OrbitData Volume/Mbitsper 7 daysDump Capabilityof S/C Avionics4176 75.168CDMU 20 411 7.2Payload 3765 67.9686.6.2 Lander Based on Liquid PropulsionTable 6.6-8: MSE to MMO data returnAs an alternative to the hard lander version of MSE (Mercury Surface Element), a soft lander version wasstudied for accommodation within the split launch version of the mission (ESA, 2000). For the price of a171


BepiColombosomewhat increased complexity and development cost, a soft lander could offer essentially the followingbenefits:• easier accommodation of geochemical payload, requiring medium-range deployment by means of a microrover,mechanically susceptible to the hard lander impact;• chemical cleanliness of the landing area, by separation of the propulsion system before impact;• possibility of extending the mission beyond the minimum required of 7 days, by including the mechanismsrequired to deploy and operate a solar power system and a heat rejection system;• possible obstacle avoidance at landing by vision-based navigation (not included in this study).The Solar Electrical Propulsion Module carries the MMO and MSE composite to the vicinity of Mercury,after which it is jettisoned. A CPM (Chemical Propulsion Module) is then required for injection of thecomposite into the highly elliptical, 400 x 12000 km, Mercury polar orbit required for the MMO operations.Soft landing requires essentially a flexible, high-performance, liquid-propellant propulsion system, able tobring the spacecraft to zero velocity, in a controlled attitude, in close proximity of its intended landing site. Thevelocity increment for injection into Mercury orbit is only of the order of 350 m/s, while the combined velocityincrement for de-orbit and landing is >4 km/s. It is therefore envisaged to combine the CPM and the propulsionsystem of MSE into in one vehicle, with a single propulsion system. The actual MSE lander (the surface station,including the payload and the power, data and communications subsystems) is a self-contained package, whichseparates from the integrated CPM in close proximity of the ground.Soft landing has been defined for this study as landing at a speed not greater than 30 m/s. With a landingsystem based on airbags (heritage from on-going European developments for the BEAGLE 2 andNETLANDER Mars lander missions), the resulting worst-case impact shock is not greater than 250 g for 38 ms.The 30 m/s impact velocity is the same as the parachute terminal velocity assumed for the above mentionedMars lander missions, allowing the same design to be re-used, and corresponds to a free fall in Mercury gravityfrom a height of 120 m above ground.Landing on Mercury shall occur in the latitude band between 84° and 86°, in either hemisphere, with norequirements on longitude from the scientific side. The near-polar location provides a milder environment than alower-latitude one from the thermal standpoint; however, it reduces considerably solar illumination of thesurface (typically up to 40% of the surface may be shadowed by local features), with consequent limitations onsolar power availability.The main design drivers for the soft lander can therefore be summarised as follows:• power and operations (including navigation during landing) require targeting an illuminated area;• propulsive braking greatly limits the useful payload mass on the surface;• soft landing requires a shock absorption device;• thermal environment constrains the landing location and the duration of surface operations;• the reference payload requires a compromise between mobility, soil penetration, and avoidance of chemicalcontamination.172Mission OverviewThe operations of the MSE mission are described in the following sequence (Figure 6.6-12):• The composite (MMO, MSE, CPM) separates from SEPM. The MSE lander is not active, and thecomposite is controlled by MMO. The composite mass at this stage is 615 kg for the 2007 launch window(on the basis of performance specified for the Soyuz/Fregat launcher and SEPM).• CPM fires (thrust 4 kN, Isp 320 s, )v = 350 m/s) and injects the composite into orbit around Mercury.• The composite loiters in a highly elliptical orbit (up to 33 days), waiting for the correct landing conditions(surface illumination).• CPM and MSE separate from MMO, which starts its scientific orbital mission. The composite mass is 439kg. MSE takes control of the composite.• CPM fires (thrust 4 kN, Isp 320 s, burn time 67 s, )v = 850 m/s), in proximity of MMO orbit periherm, andinjects MSE on a transfer orbit with a periherm of 10 km above one of the polar regions.• Approximately 75 minutes after transfer orbit injection, CPM fires again (thrust 4 kN, Isp 320 s, burn time138 s, )v = 3208 m/s), and reduces the velocity of MSE to 0 in close proximity of the ground (120 m). The(dry) mass of the composite is reduced to 99 kg.• MSE is detached from CPM and is separated by 3 small 90 N solid-fuel thrusters (2 s burn time). The CPMcrashes into the ground within 8 s of separation. The separated MSE mass is 49 kg.• The lander inflates its airbag system during its free fall.• The lander hits the ground (at about 100 m from the CPM crash site) after a 10 s ballistic flight andrebounds several times, increasing its distance from the CPM debris.• The lander comes to rest. Airbags are separated and deflate. The lander falls to the ground.


Spacecraft Composite• The lander rights itself up (if need be) and deploys power, thermal control, and communicationssubsystems.• The lander deploys the payload and starts operations.• MMO comes into visibility of the lander and a communications session is started.• Operations continue for a minimum of 7 days, and can be extended to a maximum of 35 to 70 days(according to the selected landing date).• Mercury night-time sets in, solar power is no longer available, and the MSE mission is terminated.Figure 6.6-12: MSE mission sequence of operations.This sequence of operations is assumed as baseline for the MSE soft lander version, and is based on a numberof system-level trade-offs carried out in the study and summarised in Table 6.6-9.Figure 6.6-13: Composite at launch and CPM after separation from MMO.Launch and cruise configurationFigures 6.6-13 and 6.6-14 show the CPM overall configuration and dimensions. In the preferred launchconfiguration, compatible with the standard Soyuz/Fregat fairing, a 630 mm mechanical interface with MMO isused, while any interface of CPM with SEPM is avoided. The MMO high-gain antenna is hosted inside theSEPM internal space. This configuration allows an overall mass reduction, by avoiding the need for a thrustcone between SEPM and MMO (the CPM structure is a lightweight CFRP truss and cannot transfer launch173


BepiColomboloads). The overall CPM configuration is maintained symmetrical throughout the mission. MSE isaccommodated inside the truss structure and the 630 mm MMO interface ring.PropulsionOrbital PhaseItem Baseline Second Option NoteSingle engine for orbitinjection, descent andlandingDelayed descent afterMercury orbit injectionSeparate CPMs forinjection and fordescent/landingImmediate descent afterinjectionStaging Single stage Staging by tankjettisoningBetter performance (mass andthrust), less complexityDescent trajectory aboveilluminated side of Mercuryrequired for vision-basednavigationNo major mass benefit byjettisoning, more complexityLanding System Airbags Crushable structure Airbags more compact, lighterConfiguration/CPM StructureCPM-LanderSeparationCPM on top ofMMO/truss structureLander separates fromCPM before touchdownCPM underMMO/thrust coneLander does notseparate from CPMTable 6.6-9: MSE system-level trade-offsMass saving, but operationalconstraints on MMOAvoidance of chemicalcontamination of landing site,lighter/simpler landing systemThe chemical propulsion subsystem uses a single 4 kN main engine, mounted on the CPM centreline underthe truss. Four propellant tanks (2 NTO, 2 MMH) and one helium pressurant tank are carried by the modulestructure.Thermal control of CPM before separation from MMO is insured by MLI insulation wrapped around CPM,with titanium foil patches near the main and RCS thrusters for protection against plume heating. Heater power issupplied from SEPM and/or MMO if required.Figure 6.6-14: Composite at launch and after separation from SEPM (overall dimensions in mm).An alternative configuration (Figure 6.6-15) was considered in the study, to comply with the requirement todeploy the MMO high-gain antenna in support of communications during the cruise to Mercury. In this option,CPM is partially hosted inside SEPM, while a thrust cone (sandwich structure with CFRP facesheets) is used tointerface with MMO. The propulsion system still has the same 4 propellant tanks, but 2 pressurant tanks (allequatorially mounted at the same level), because of the need to reduce the overall height of CPM. In this option,the structural mass is increased by about 10 kg.174


Spacecraft CompositeSeparation and DescentAfter release from MMO, CPM is de-spun and 3-axis stabilised. The AOCS for the autonomous flight ofCPM is based on:• a sensor suite: star tracker, inertial measurement system (3 gyroscopes, 3 accelerometers), range sensingsystem;• actuators: main engine, 4 RCS engines, propulsion drive electronics.Figure 6.6-15: Alternative composite configuration at launch.The proposed range sensing system is based on an optical device, tracking features on the ground from oneimage to the next. This tracking requires modest onboard resources (compared with a radar altimeter), becausethe inertial measurement system will provide information of the vehicle motion between images, so that afeature can be searched for within a small window of uncertainty. The relative motion of tracked points willreveal the vehicle altitude, knowing the distance between the points from which the images were taken. Thisdistance is not directly known, due to dispersions in onboard knowledge of the velocity, but, when a measuredacceleration is applied to the vehicle, the known change in velocity can be used to scale the distance to theground. Notably, several ground features distributed over the image, depending on the local lighting conditionsand surface morphology, are tracked (notice that this requires that approach to landing occurs over a sunlitportion of the Mercury surface). RSS filtering yields an estimate of the average ground plane in the image.The main engine was selected on the basis of the following criteria:• high thrust required for descent optimisation (near-optimal in the range 1.5 to 4 kN, 4 kN selected on thebasis of commercial availability);• high Isp required, for mass saving (315 s advertised by manufacturer, improvement to 320 s envisaged bythis study, by nozzle and combustion chamber modifications);• re-startability (3 burns) required to cover the whole injection, descent and landing sequence;• no thrust level modulation capability required (the robust landing system, based on airbags, does not requirea precision-landing capability).The consequent high thrust-to-mass ratio results in a highly dynamic system. Reaction control isimplemented by four 20 N MMH/NTO thrusters, which are a reasonable compromise between the high thrustlevel required for torque compensation, and the low thrust level required to provide small minimum-impulse bitsduring the orbital phases. On the other hand, the short main engine burn times relax requirements on gyroscopedrift performance.The low power levels and number of cycles suggest the use of Li-Ion batteries, to reduce mass and thermalcontrol issues. In principle, only a primary battery is required for descent and landing operations (up to 1.5 h intotal), but a rechargeable battery is preferred (low extra mass, and possibility to top up battery charge from theMMO solar array after >2 years cruise); capacity is 6.3 Ah. Thermal control of the lander from MMO separation(lander avionics activation) to landing (lid opening) relies on thermal inertia.A descent imager (CLAM-D) is part of the reference payload, with the purpose of documenting the approachand landing, and to characterise the structure of the surface. It could take, indicatively, 10 images during thedescent, the last one as low as 100 m above the surface (just before MSE lander separation from the spentCPM). No communications from the MSE are envisaged during descent and landing: MMO will not be visible175


BepiColomboduring landing and direct communications from MSE to Earth would have a severe impact on the system.Images from CLAM-D are compressed and stored in MSE for later downloading. The AOCS star sensor andrange sensor, and CLAM-D are mounted on CPM, to ensure adequate fields of view. The transfer orbit profileand the final descent are shown in Figure 6.6-16.MeSE Soft Lander 4000 N 2-Burn Descent: Transfer and Final Descent300012Altitude [km]20001010008Planetocentric z [km]0-100064-20002-300008584838281807978777675-4000Latitude [deg]-3000 -2000 -1000 0 1000 2000 3000Planetocentric x [km]Figure 6.6-16: Transfer orbit and final descent altitude profile.LandingEjection of the lander from CPM is initiated by 3 kick-off springs and is carried out by three 90 N separationthrusters, with 3 axial rails to guide the lander out of CPM during the operation. The landing system consists of3 inflatable airbag envelopes tightened together by a rope into a spherical shape, 2.3 m in diameter, with a totalvolume of 6.4 m 3 (stowed: 0.0113 m 3 ). The envelope material is a 0.2 mm thick rubber-coated aramid fabric; itis inflated by 3 pyrotechnic gas generators (>95% N 2 ) in about 1 s. The airbag configuration is shown in Figure6.6-17. At a touchdown velocity of 30 m/s, the resulting worst-case impact shock is 250 g for 38 ms. When thetightening rope is cut after landing, the 3 envelopes spring back to their natural spherical shape because of theirinternal pressure, and separate from the lander, which falls to the ground from a height of about 1 m.Airbags attached to lander(one of 3 airbag envelopes not shown)Separation of airbags from landerFigure 6.6-17: Airbag Landing System.176


Spacecraft CompositeSystem accommodationThe MSE lander has a cylindrical body (600 x 300 mm) with a rounded side, to ensure that it will come torest on either base. An autonomous attitude recovery capability after landing is ensured by the deployable lid.The lander body has a sandwich construction with CFRP facesheets; metallic facesheets may be considered, toimprove radiation protection with a limited mass penalty. The lander body houses all instruments and ancillaryequipment and mechanisms, and provides them with thermal protection; equipment accommodation inside thelander is shown in Figure 6.6-18. The equipment to be deployed is stowed in a gap between the lid (solar panel)and the radiator. Only the micro-rover is stowed in a dedicated compartment under the radiator, with its ownexit opening (protected by a flexible sunshade).CameraElectronics3 ReleaseThrusters3 GasGeneratorsCommunicationsElectronicsPowerModuleData Handling Electronics(under Microrover)HP 3 MolePayload Electronics& BatteriesFigure 6.6-18: Equipment accommodation inside the lander.The solar array is deployed on a mast (total height 1.15 m above ground). Pointing and power transfer isperformed by a solar array drive at the base of the mast. Array pointing is required to be offset at an angle of upto 70° away from the Sun direction, for thermal reasons (maximum temperature 130 C); 50% of the arraysurface is covered with OSRs; Sun acquisition is done by solar array current and temperature sensors reading;GaAs solar cells are selected for higher efficiency. As an option, a primary battery could be added for landersurvival until first MMO visibility; in case of landing in a dark area, a 2-day lifetime is possible with Li-Thionyl-Chloride (Li/SOCl 2 ) batteries (2 kg, 300 Wh/kg). A fully regulated power bus at 28 V is available forall standard equipment.Thermal control relies on heavy insulation, with MLI blankets wrapped around the whole lander body. Allinstruments inside the lander body are thermally coupled to a zenith-facing 0.16 m 2 OSR radiator, capable ofrejecting 50 W. The radiator is protected from sunlight by a deployable sunshade (0.21 m 2 , 50 :m-thick, whitepaintedtitanium foil), hanging from the edge of the solar array, and rotating with it to follow the sun. Thesunshade is unfolded and straightened by pre-stressed tapes.Cross-dipole antennaSunshadeLid (solar array)CLAM-SHP3 mole(inside body)MLMAG & SEISMOAXS micro-rover(tethered)Micro-rovercompartment doorFigure 6.6-19: Deployed configuration on surface.177


BepiColomboA UHF data relay (via MMO) is considered as a baseline, with mass and power requirements of 2 kg and


Spacecraft Composite6.6.3 Concluding Remarks on the Surface ElementThe current study results demonstrate the flexibility of options available for the design of MSE. The hardlander(Section 6.6.1) baseline has been retained for the estimation of the system budgets. However numerousoptions have been generated from a parallel study activity (Section 6.6.2), which are worth being considered forincorporation at a later stage of the design evolution. Major items are summarised in Table 6.6-12.In particular, the option of using liquid propulsion only is of great interest, since it has the potential ofreducing the overall mass at system level (Section 6.1.3). This, coupled with a more capable GNC, would allowthe adoption of a lower landing speed and of an airbag landing system, more suitable for the operation ofgeochemical payload in a chemically uncontaminated environment. Therefore these options should be given dueattention in the future design evolution.On the other hand, extended life time on the surface may well result too ambitious within the overall missioncontext, especially considering the risk induced by the significant portion (40%) of the landing area withoutsolar illumination, and by the long-term operation of mechanisms in a harsh thermal environment.179


BepiColomboItem Baseline Option RemarksPropulsion forfinal brakingLanding speed(engine shutoffaltitude)ImpactattenuationsystemLife time onsurface (powersystem)Solid propellant86 m/s(1 km)Propellant tankused ascrushablestructureLimited to 7days(batteries only,total of 473 Wh)Liquidpropellant30 m/s(120 m)• A liquid-propellant engine gives better massperformance at the expense of higher cost andcomplexity• Integration of MSE engine with CPM would bebeneficial for system-level mass margin• Lower engine shut-off altitude requires morecomplex GNC (e.g. based on inertial measurementunit and optical range sensor)• Higher landing speed is generally more suitablefor subsurface physical properties measurements(penetrator-based)• Lower landing speed is generally more suitablefor surface geochemical measurements (requiringmechanisms for surface deployment and mobility)• Lower landing speed makes synergies with thedesign of European Mars landers possibleAirbags • Crushable structure is more sensitive to the natureof terrain and to the impact angleUp to 70 days(up to 30 Wcontinuousaverage powerfrom solararray)• Use of propellant tank as crushable structure willinduce chemical contamination of the landing site• Airbags are a very robust system, but will tend toend up in a terrain depression• Infrequent contact with MMO relay is possibleonly 18 times during 7 days at 9-hour intervals,limiting the flexibility of payload surfaceoperations (i.e. re-targeting of micro-rover)• Solar array-based system may be hampered bylocal illumination conditions (low sun elevation inpolar region)• Solar array and sunshade deployment and longtermoperation require mechanisms, representingsingle-point failures on the MSE mission criticalpath• Limited solar-array capability may be added to thebaseline (7-day) configurationData volumereturned frompayloadUp to 68 Mbover 7 days(batteries only)Gb levelachievable over70 days (solararray poweravailable)• Availability of energy is the main limiting factorfor the data volume from MSE• Large data volume from MSE has impact onMMO operations (MMO payload may have towork at reduced capacity for the duration of MSEpayload operations)• CLAM-D ideally requires a large data volume tobe downloaded after landing (in the order of 50-200 Mb)Table 6.6-12: MSE baseline and options180

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