Exp AstronDOI 10.1007/s10686-010-9209-y

REVIEW ARTICLE

Mission-constrained design drivers and technicalsolutions for the MAGIA satellite

Giorgio Perrotta · M. Stipa · D. Silvi ·S. Coltellacci · G. Curti · G. Colonna ·T. Formica · V. Casali · T. Fossati ·F. Di Matteo · M. Zelli · M. Rinaldi ·L. Ansalone · A. Di Salvo

Received: 8 April 2010 / Accepted: 23 November 2010© Springer Science+Business Media B.V. 2010

Abstract The Mission MAGIA (Missione Altimetrica Geofisica GeochImicalunAre) was proposed in the framework of the “Bando per Piccole Missioni”of ASI (Italian Space Agency) in 2007. The mission was selected for a phaseA study by ASI on February 7th 2008. The tight budget allocation, combinedwith quite ambitious scientific objectives, set challenging requirements for thesatellite design. The paper gives a fast overview of the payloads complementand of the mission-constrained design drivers, including cost minimization, riskreduction, and AIT flexibility. The spacecraft architecture is then outlined,along with an overview of the key subsystems and trade-offs. Some detailsare given of a Moon gravitometric experiment based on a mother–daughtersatellite configuration with the daughter being a subsatellite released fromthe MAGIA satellite and intended to circle the Moon at a very low altitude.Budgets are appended at the end of the paper showing the key study results.

Keywords MAGIA · Lunar orbiter · Satellite design · Subsatellite ·Moon orbiter · Spacecraft subsystems · Spacecraft modeling ·Satellite trade-offs · Propulsion design · Thermo-structural design ·Power subsystem design · Optical experiments accommodation ·RF experiments · Particle experiments accommodation · Moon-orbiter toEarth communications · Ranging · Gravitometric experiment

G. Perrotta (B) · M. Stipa · D. Silvi · S. Coltellacci · G. Curti · G. Colonna · T. Formica ·V. Casali · T. Fossati · F. Di Matteo · M. Zelli · M. Rinaldi · L. AnsaloneSpaceSys, Via Latina 293, Rome, Italye-mail: giorgio.perrotta@spacesys.eu

A. Di SalvoRheinmetall Italia, Rome, Italy

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Fig. 1 Key elements of the MAGIA Mission

1 MAGIA mission

The MAGIA mission, whose key elements are shown in Fig. 1, has thefollowing scientific objectives:

– Detailed study of the internal structure of the Moon through its gravity;– Study of the polar and subpolar regions in terms of their morphology and

mineralogy;– Study of the lunar exosphere and radioactive environment;

The mission intended also to contribute to the fundamental physics via mea-surements of the gravitational redshift, and to perform test in view of thesecond-generation lunar laser ranging.

The scientific mission was implemented with a suite of instruments andexperiments as shown in Table 1.

A total payload mass plus control electronics, thermal control and harnessof less than 60 kg results, which is fairly credible since the most massive units

Table 1 Payload complement to accomplish the scientific mission

Instrument Acronym Mass (kg)

Spectrometer context camera CAM-SIR 11High resolution camera CARISMA 4Radar altimeter and radiometer RAR 9Gravitometric experiment dual accelerometer ISA 6.1 (+6.1 in subsatellite)Neutral particle detector ALENA 1Particles spectrometer RADIO 0.3 (est.)CCR array VESPUCCI 3CCR-MoonLight MoonLight-P 1.2Radio science Radio science X- and S-band RF links

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Table 2 Orbit parameters forthe lunar mapping orbit

Lunar mapping orbit

Semi-major axis 1,838 kmEccentricity 0.00675Inclination 89.99Argument of perigee 270◦Orbital period 2 h

are already existing or are derived by similarity to already-flown ones. Withthe exception of the Radar Altimeter and Radiometer and the MAGIA-borneaccelerometer (the other accelerometer is installed on board the releaseablesubsatellite) all instruments are housed in a thermally controlled modulemechanically decoupled from the rest of the spacecraft.

The Earth–Moon transfer is a dimensioning space segment factor due tothe propulsion requirements. Different lunar transfers were analyzed includingHohmann-like and variants of the Weak Stability Boundary trajectories. AHohmann transfer was selected for its simplicity and short time duration.The two launchers taken into consideration, as requested by the Customer,are Vega and Soyuz/Fregat, which represent European small and mediumclass options in terms of capability and cost. The Soyuz/Fregat launcherresulted to be fully compatible with the mission and transfer requirements andwas therefore chosen, allowing a direct lunar injection, while minimizing thepropellant mass to be embarked.

A circularization maneuver around the Moon, to insert the spacecraft intothe operational orbit, is required by lowering the orbit altitude with respect tothe arrival conditions. A polar frozen low lunar orbit (LLO) was selected fora preliminary characterization from an operational point of view (e.g., lunarcoverage, ground visibility etc.); the orbital parameters are shown in Table 2.

An uncontrolled LLO, leaving the orbital plane essentially unchanged inan inertial frame, allows in principle a complete coverage of the lunar surfaceduring a sidereal month. The Moon advances about 15◦/day along its orbit, andwithin this time span, a satellite injected on a 100 km orbit completes about 12orbits around the Moon. As a consequence, the angular separation betweentwo subsequent orbits in a frame rotating with the Moon is of the order of 1◦.

The maximum duration of spacecraft routine solar eclipses lasts about45 min, roughly corresponding to 40% of the orbital period; total lunar eclipsesrepresent the worst-case period for the spacecraft without sunlight.

Two distinct mission phases are foreseen: the first devoted to lunar mappingand imaging and the second to the gravity experiment. Two slightly different

Table 3 Orbital parametersfor the gravity experiment

Gravimetric experiment orbit

Semi-major axis 1,798 kmEccentricity 0.00675Inclination 93.00Argument of perigee 270◦Orbital period 2 h

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Table 4 Earth–Moontransfer velocity incrementsrequirement

Maneuver (s) �V (km/s)

LOI (Hohmann transfer) 0.8Lunar imaging phase orbital control 0.070Orbital transfer 0.093Safety margin (+5%) 0.050Total 1.012

nominal orbits have been respectively selected for the two phases. For the6 months of mapping and imaging, it is highly desirable to have a polarorbit, since it guarantees the best coverage of the entire lunar surface. Theexperiment requires that the satellite altitude be maintained within a rangeof 100 ± 30 km. Therefore, orbital correction maneuvers are preliminaryplanned every 40 days to restore the eccentricity to its nominal value. Since theduration of this mission phase is fixed at 6 months, five correction maneuversare needed. Each maneuver consists of two burns for a total of 14 m/s. Thetotal Delta-V necessary for maintaining the orbit inside the nominal box forthis phase is then 70 m/s.

At the end of the lunar mapping phase, the spacecraft must be transferredto the gravimetric nominal orbit (Table 3), by changing orbit semi-major axisand inclination, required to cope with scientific experiment requirements. Thelatter maneuver, to achieve an inclination change of about 3◦, is the mostexpensive, requiring a Delta-V of about 85 m/s; a total Delta-V of 93 m/sis estimated for the overall orbital transfer, including the semi-major axistrimming.

The orbit control strategy during the gravity experiment phase is compli-cated by the presence of a releasable subsatellite, which has no maneuveringcapability. The gravitometric experiment requires at least 1 month of opera-tions, at a mean radius of 1,798 km, though a longer mission duration of up to3 months was considered.

The total velocity increment for the overall mission that must be providedby the spacecraft propulsion system, is summarized in Table 4. Let us stressthat the first maneuver to inject the spacecraft into the Lunar Transfer Orbitis assumed to be performed by the launcher. To accomplish all complexmission maneuvers, a double propulsion system was baselined, including botha hydrazine system for spacecraft orbit transfer and reorientation, and a cold-gas one for attitude control tasks.

2 Spacecraft design approach

The spacecraft architecture consists of three functionally, physically and tech-nologically independent assemblies:

Propulsion module; Platform (or Service) module; Payload module

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The sum of these three modules can be enveloped by a parallelepiped withdimensions: 700 × 1,400 × 2,050 mm.

The propulsion module hosts both a hydrazine-based main propulsionsystem and a cold-gas based auxiliary propulsion system. It also hosts thesubsatellite along with its release system.

The platform (or Service) module hosts the majority of the electronicsubsystems, with the exception of the star sensors which are integrated inthe payload module and of the X-band high-gain antenna (HGA) which isaccommodated on the upper part of the payload module as well.

The payload module accommodates most payloads—with the exception ofthe radar payload, a high sensitivity accelerometer (ISA), and a high stabilityclock. The payload module has been carefully designed to meet the tightoperating temperature limits of the payload equipments.

The spacecraft design was conceived to be largely built around previouslyqualified items, when available, or on COTS further subjected to delta-testswhen their heritage or status did not meet minimum quality levels consideredadequate within the cost limitations of the Programme.

A prototype approach was also baselined, since most units were derivedfrom already-flown ones, with the exception of new instruments that had to besubjected to a qualification campaign with somewhat reduced severity levels,however.

3 Spacecraft structure

The spacecraft structure encompasses all three modules using different tech-nologies matched to the operational requirements of each. More specifically,the propulsion module, shown in Fig. 2, is an open-truss structure of carbon-

Fig. 2 Propulsion module,tanks and subsatellite

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Fig. 3 Service module,interior

fiber-reinforced plastic (CFRP) tubular elements connected with high-strengthaluminum alloy joints.

It carries the tanks of the propulsion system (mono-methyl hydrazine fueland a pressurant) and the Attitude Orbit Control Subsystem thrusters, thepropulsion system harness (valves, tubes etc.) and the subsatellite deployer (aclosed, box-shaped, aluminum alloy sandwich structure carrying the subsatel-lite for the gravitometric experiment).

The propulsion module is connected to the service module by high-strengthscrews permitting a smooth transfer of loads. The service module, shownin Fig. 3, is a closed aluminum alloy sandwich structure with high-strengthaluminum frame carrying longitudinal and lateral loads. The lateral sand-wich panels are integrated with aluminum-machined plates with stiffeners toadequately support the electronic boxes accommodated on the panel whichcontribute to the overall satellite stiffness. The interface with the propulsionmodule and the top panel—which interfaces with the payload module—arethick core sandwich panels to meet the requirements of dimensional stability.The Service module carries most of the electronic boxes including the Radar

Fig. 4 Payload module,interior

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Fig. 5 MAGIA spacecraftstructure in three modules:launch configuration

payload; and—along with the propulsion module structure—supports the largelightweight sandwich solar panels via a truss structure made of small-sectionCFRP tubular elements. The payload module, illustrated in Fig. 4, is a closedbox made by sandwich structural panels carrying optical payloads and startrackers. The base panel is a dedicated optical bench in order to have acommon stiff interface and to permit the correct alignment between differentinstruments optics and star sensors.

The mechanical connection with the service module is designed in order toreduce the thermo-elastic deformations coming from the orbital thermal loads.

The MAGIA satellite, in both its launch and in-orbit configurations, isshown in Figs. 5 and 6. A detailed Finite Element Model was developed toperform both static and dynamic analyses of each module and of the fullspacecraft. A modal frequencies analysis was performed along the X-axis, asthe most representative, to obtain modal participation mass factors. The resultsindicate compliance with the launcher (Fregat/Soyuz) frequency (>20 Hz onlateral axes). Besides, a rapid overview of the modal analysis enabled topinpoint that the lower side of the elements of the propulsion module truss arethe more stressed elements. Detailed static analyses were then performed on

Fig. 6 MAGIA spacecraft:in-orbit configuration withtwo solar panels deployed

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each module to better assess the structure criticalities. Several areas requiringlocal improvements were thus identified and characterized.

4 Propulsion subsystem

Three design alternatives were considered:

(a) Independent hydrazine and cold-gas (N2) propulsion systems. The pri-mary propulsion system operates in a pressurized mode, requiring ahelium-filled pressurant tank; and a separate N2 auxiliary propulsionsystem operating in a blow-down mode, the N2 being contained, at highpressure, in a third tank;

(b) A hydrazine/nitrogen propulsion system with two interconnected tanks,one containing hydrazine and the other compressed nitrogen acting asa pressurant for the hydrazine system while operating in a blow-downmode for the cold-gas, auxiliary, propulsion system;

(c) An all-hydrazine propulsion system operating in a blow-down mode forboth the primary and auxiliary propulsion system.

Conservatively, the first configuration (a), shown in Fig. 7, with three tanksand two independent propulsion systems, was selected as a baseline, whichis largely based on commercial, space-qualified components: tanks from

N2

fill/drainvalve

pressuretransducer

filter

pressureregulator

thrusters thrusters

pressuretransducer

reliefvalve

Fig. 7 Pressurized hydrazine for orbit control (on the right); and cold nitrogen propulsion system,operating in a blow-down mode, (on the left) for attitude control

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PSI/ATK or ARDE’, hydrazine thrusters from EADS, valves and cold-gasthrusters from Moog or Encore.

Since the orbital maneuvers require repointing and will be implementedpiecewise, a commercial, qualified, thrust motor (EADS CHT 20) providing20 N thrust is employed. About 140 kg of hydrazine will be stored in a quitelarge elliptical tank (type 80420-1 from ATK/PSI). A pressurized, helium-based, pressurant tank is used to keep high the motor Isp even when theHydrazine tank will be almost depleted. The pressurant is contained in atitanium spherical tank (type 80202-1 from ATK/PSI) also hosted in thePropulsion module.

The propellant tankage will be capable of providing the velocity incrementof Table 3 for a satellite dry mass of 232 kg, which represented a target massfor the feasibility study phase.

The cold-gas propulsion system is used for reaction wheels desaturation, aswell as to support spacecraft reorientation pre-post hydrazine engine firing.The N2 tank (a PSI/ATK type 80345-1) is sized to accommodate 5 kg of N2at high pressure, and will operate in a blow-down mode. The cold-gas space-qualified commercial thrusters will be in number of 8, in two quadruplets offour thrusters each. The thrusters will be characterized by a thrust in the 20- to50-mN range and an Isp around 50 to 60 s. The thruster can be from Moog orother space-qualified microthrusters suppliers.

The computed dry mass of the propulsion subsystems amounts to 33 kg,excluding the structure supporting the various components.

5 Power subsystem

The configuration of the solar array is quite unconventional due to severalreasons: the time evolution of the satellite orbit with respect to the Sun; thenadir-pointing satellite attitude in a Moon orbit; the payloads duty cycles;and the requirement for an unobstructed field of view for heat dissipation

Fig. 8 Solar array geometry

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Fig. 9 Solar array concept

by radiation. The solar array dimensions result from the combination of theabove factors. During the study, several solar array geometries were analyzed,starting from that schematically shown in Fig. 8 which gave rise to many othervariants, the last of which is conceptually shown in Fig. 9.

The rationale for the configuration shown in Fig. 9 is the need for rejectingheat from the sides of the platform module which is densely populated withpower dissipating electronic units while retaining the total solar array area andshape of the initial configuration.

During launch, the in-orbit deployable panels will be restrained by means ofreleasable double clamps. Accordingly, there will be six identical solar panelseach including 22 strings in parallel of 10 cells in series each, for a nominalvoltage around 26 V at Tamb (1,320 cells in total).

The solar array is thus composed of four body-fixed panels—two on thedorsal side and two lateral panels, canted at 22◦w.r.t. the vertical—and twoin-orbit deployable panels. The fixed dorsal and lateral panels serve also toblock the Sun rays from reaching the propulsion tanks; while the two in-orbitdeployable panels, also canted by about 22◦ with respect to the local vertical,are deployed to allow the platform module sides to radiate heat towards freespace.

High-efficiency, triple-junction, GaAs cells with an efficiency of 28% willbe used, with dimension of 80 × 40 mm and a space of 2 mm between any twoadjacent cells. A lower cell efficiency of 26%—leading to a considerable costsaving—might become feasible by reducing the satellite and payloads powerdemand. In the baseline each string of the Solar Array will have 10 cells (26Voc at Tamb and 31.5 Voc at −75◦C, BOL) and 22 strings in parallel. In total,there will be six panels all identical with dimensions of 924 × 820 mm.

A simulator has been developed in Matlab to analyze, as indicated inFig. 10, the generation of electric power from solar panels depending on the

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Fig. 10 A Matlab simulatedanalysis of the generation ofelectric power from solarpanels

position along the Moon orbit and the Sun vector incidence angle. The powerconsumption along the orbit is also simulated; therefore the charge–dischargeof the battery is taken into account. The following relevant parameters werecomputed:

– Average power generation = 622.6 W– Maximum power generation = 933.9 (lat 0◦/long 0◦)

The battery is based on a Li-ion technology which has achieved a fairmaturity stage. The battery is sized for a nameplate capacity of 1,000 Wh;it is planned to drain energy from the battery up to a d.o.d. between 20%and 25%, although the technology could well support d.o.d larger than thisthroughout the about 3,300 charge–discharge cycles corresponding to a nom-inal 9 months mission duration. The Electric Power management implementsa semi-regulated bus architecture. In sunlight, a peak power tracking logic isactive to better adapt the variable load to the voltage–current characteristicsof the solar array. The battery charge circuit adapts the semi-regulated bus tothe voltage–current regime required by the Li-ion battery. In eclipse the bus isfed by the unregulated voltage from the discharging battery.

6 Thermal control subsystem

The first priority was to optimize the spacecraft thermal control during thenominal Moon mission. Several iterations were performed: the approachfinally adopted was also capable of supporting the spacecraft operations inits parking orbit and throughout the Earth–Moon transfer phase. The basicapproach was to strive to achieve average Payload module temperatures lower

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Fig. 11 View of body-fixedand in-orbit deployed solarpanels vs lateral Servicemodule surfaces

than those of the Service module, while keeping the propulsion subsystemcomponents well within their safe operating temperature range.

To achieve this result we have maximized the radiative surfaces of thepayload module and minimized the thermal exchange between payload andplatform modules. To maintain the Service module temperatures within thehosted equipments allowed temperatures, we have chosen to in-orbit deploytwo of the six solar panels in order to maximize the radiation into free space ofthe two side panels of the Service module (panels 8 and 9 of Figs. 11 and 12).

Besides, we have positioned four of the six solar panels in such a way toshield the propulsion components (mainly the hydrazine, He and N2 tanks)from the Sun rays. All tanks were also wrapped with multi-layer insulation(MLI). Two of the four solar array panels used to shield the propulsion

Fig. 12 View of tanks of thePropulsion module vs Servicemodule and solar panels(fixed dorsal and deployedlateral)

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Fig. 13 Units inside theService Module

module from the Sun rays are canted by 22◦ in order both to improve theirenergy collecting efficiency as well as to improve their heat radiation capabilitytowards free space, at least during portions of the Moon orbit.

In Figs. 13 and 14, the radiating surfaces are green-colored, the contact areasare in red. To improve the thermal conductivity we used a thermal filler CHO-THERM 1671 throughout. To thermally isolate units from the baseplate, weused thermal washers with low thermal conductivity values.

This first optimization resulted in absorptivity/emissivity values pictoriallyshown in Figs. 15, 16, and 17. We have also used four thermal straps (AmocoP100 Carbon fiber. Radius = 5 mm, length = 20 mm) to connect solar panelY+ to service module faces Z+ and Z− (for the axes see Fig. 11). The strapsare placed inside the beams that link the service module to the solar panel.

It was indeed necessary to lower the temperature of Service module tobring a few critical units (e.g., the high sensitivity accelerometer) within theirspecified operating temperature range. Besides, we found necessary to lowerthe temperature of the face Z+, facing the payload face Z−, to reduce the heatflux. We thus achieved a doubly positive result: a strong reduction of the solarpanel gradients and of the heat flux towards the payload module.

Fig. 14 Units inside thePayload module

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Fig. 15 Absorptivity α

Fig. 16 Emissivity ε

Fig. 17 α/ε ratio

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Table 5 Coatings andmaterials eventually used forMAGIA thermal controloptimization

External surface Internal surface

Payload:FACE X+ (5) OSR 146434 Black PaintFACE X− (2) OSR 146434 Black PaintFACE Y+ (3) OSR 146434 Black PaintFACE Y− (4) MLI 20 layers Black PaintFACE Z+ (1) OSR 146434 Black PaintFACE Z− (6) MLI 10 layers Black Paint

Propulsion:Tank Hydrazine (20) MLI 10 layersTank He (21) MLI 10 layersTank N2 (22) MLI 10 layers4 long beam (18) Black Paint Black Paint8 short beam (19) Black Paint Black Paint

Platform:FACE X+ (9) White Paint Black PaintFACE X− (8) White Paint Black PaintFACE Y+ (10) White Paint Black PaintFACE Y− (12) MLI 10 layers Black PaintFACE Z+ (7) Kapton, aluminized Black PaintFACE Z− (11) White Paint Black PaintSolar Y (13) Solar cell Black PaintSolar X+ (14,16) Solar cell Black PaintSolar X− (15,17) Solar cell Black Paint

The same strategy was followed to meet the propulsion tanks temperaturerequirements. Indeed, we used two straps (length = 40 mm, radius = 5 mm)that, like the previous ones, are placed inside the beam linking the solar panelto the propulsion module structure.

As a result of extensive trade-offs, materials and coatings used for internaland external surfaces are shown in Table 5. The simulation results, relevantto the Moon-orbiting phase, show, in Figs. 18 and 19, that all equipmentremain within the operating or storage temperature ranges, depending on theiroperative status. However, to achieve good results, it was necessary to useheaters in a selective way, consuming between 5 and 30 W DC power fromthe spacecraft bus.

Fig. 18 External panels temperatures with Sun rays in the Y–Z plane

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Fig. 19 External panels temperatures with the Sun rays hitting the spacecraft laterally (i.e, in theX–Z plane)

Besides, very preliminary analyses showed that, during the transfer orbit,the temperatures go to the lower limits. This might require an increase of theDC power drain by heaters or else a further optimization of materials andcoatings to also take into account the detail requirements of the few equipmentthat must operate also during the transfer orbit.

7 Attitude and orbit determination and control

The satellite must determine its attitude in three axes with an accuracy notworse than: 0.01◦ at 3 sigma. The following sensors are envisaged:

– High accuracy star sensors: two pointing towards different directions. Theintrinsic accuracy of the star sensor is of the order of: 0.01◦

– Sun sensors of medium accuracy. These sensor are mainly used in the safemode and, if cooperating with the GPS receiver, supports the orbit resti-tution on board as an aid to implement autonomous or semi-autonomousnavigation;

– Three-axis rate gyros. These serve to stabilize the desired pointing direc-tion against disturbances

The spacecraft attitude must be controlled with an accuracy better than 0.1◦using the following actuators:

– Reaction wheels in classical tetrahedral arrangement. Wheels, with a nom-inal torque capability of 20 mNm, will perform small but nearly continuousattitude adjustments around the c.o.g.

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– Cold-gas thrusters: used to perform wheels desaturation. Two quadrupletsof thrusters are used, in the plane passing through the satellite c.o.m.

Orbit control will be done using the hydrazine propulsion system. The mainengine thrust vector orientation before thruster firing will be controlled usingthe attitude control reaction wheels, with the cold-gas propulsion system actingin back-up.

Precise orbit determination will be done on ground utilizing the range andrange-rate measurements at X-band performed in the context of a RadioScience experiment. However, as a back-up, the satellite orbit determinationcan be estimated on-board using an orbital propagator suitably initialized andupdated.

Furthermore, the satellite will carry an experiment aiming at assessing theusability of the GPS signals at very large distances from Earth for spacecraftlocalization and navigation purposes.

A survey of sensors and actuators was performed and key results arereported here below:

– Sun sensors: the AeroAstro coarse Sun Sensors seem a good choice for thelost-in-orbit and safe modes: these sensors are small, have a 120◦ full anglecircular f.o.v; ±5◦ accuracy; no power drain; 10 g each. At least five units,maybe six, should be installed to guarantee the Sun visibility by at leastone sensor in case of loss of pointing.

– Star sensors: among the several devices present on the space market, theAeroAstro miniature Star Tracker seems the most interesting providing a±70 arcsec accuracy (±0.02◦), at a rate of 10◦/s at an update rate of 1 Hz; acatalogue of 600 stars of fourth-order magnitude; a power consumption of2 W, a mass of 0.42 kg.

– Rate gyro: among all companies surveyed, two are the most interesting:Fizoptika, and Systron–Donner. Fizoptika offers a FOG type FG 035Qwith excellent performance: a 250 g per axis; a 1 W power drain, and a0.1◦/hour bias stability; and a measurement range up to 100◦/s. Three suchdevices will consume 3 W, with a mass of 0.75 kg, and a superb performanceenabling to maintain the satellite attitude using an orbit propagator and thegyro package. Thus, the need for a second star sensor would be confined tothe operation of the high resolution camera. The type QRS11 of Systron–Donner is a Coriolis-based space-qualified rate gyro with a weight of 60 g,0.4 W power drain; a measurement range of up to 100◦/s. However, it hasa short-term stability of 0.01◦/s (36◦/h) which is too much to support openloop attitude control in eclipse without the aid of a star sensor and reactionwheels;

– Reaction wheels: among all companies surveyed the SSTL microwheel10SP-M seems a good alternative, with a 1 kg mass, 5 W power drain atpeak torque and 0.7 W at constant speed. The 0.42 Nms momentum will,however, require more frequent desaturations than initially envisaged.

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8 The subsatellite

Unlike previous gravitometric experiments, based on very simple subsatellitesin free fall around the Moon, MAGIA aims at achieving a much better resolu-tion of the Moon gravity field. To this end, MAGIA envisages two extremelysensitive accelerometers, one on board the mother satellite and one insidethe releasable subsatellite, whose purpose is the removal of non gravitationalexternal forces acting on the bodies. The local gravity is measured by theDoppler change, in an S-band subsatellite-to-mother-satellite l.o.s radiolink.The S-band radiolink will have a low-power transmitter on board the subsatel-lite and a receiver on board the mother satellite. Since the subsatellite will nothave any device for attitude control, the S-band antenna will be essentiallyomnidirectional, implemented with six crossed dipoles. The accelerometerputs severe requirements on the attitude knowledge, power consumption andthermal control of the subsatellite, therefore its design constitutes a “projectwithin the project” which was only initially tackled.

The cube-like subsatellite will be housed in a box-shaped parallelepipedlocated in the lower part of the propulsion module. A lid will protect it duringorbit transfer and the first three months of operation in the nominal Moonorbit and will be opened before the subsatellite release which will be done bysliding it on two pairs of rails. The subsatellite power subsystem is a criticalissue. All six sides of the cube will be covered with GaAs high efficiency solarcells: the baseline dimensions, about 0.1 m2 area per panel, will produce—atnormal incidence in sunlight—about 27 W, at a bus voltage around 16 V. A Li-ion battery is foreseen (e.g. type Sanyo 18650) in a 4S3P configuration (12 cellsin total) to achieve a less-than 30% d.o.d. over one half of the orbit period.At 12 orbits/day and 6 months of projected maximum lifetime the batterywould be subjected to 2,160 charge–discharge cycles, which is a fairly modestrequirements for today’s Li-ion batteries. A Direct Energy Transfer bus wasbaselined due to its simplicity and effectiveness.

The subsatellite thermal control must restrain the temperature differentialbetween the subsatellite outer skin and the Accelerometer for which a targettemperature interface range of T-amb ±12◦ is aimed at. To obtain this result,the cube-like body must carry a white-painted area, through which to radiatethe dissipated heat. We have used four carbon straps (5 cm × 11 cm) toimprove the heat conduction from the accelerometer to the radiating strips.Besides, all other internal components are isolated from the cube-like body bymeans of thermal washers. A MLI with 5 layer is placed inside the body tofurther thermally decouple the electronics from the subsatellite outer surfaces.To improve the thermal situation during eclipses, we will use four heaters. Theanalyses have shown that with this design, the accelerometer remains in therange of ±12◦C, while the solar cells stay in a range of +110/−90◦C.

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9 Communications

Besides the conventional S-band communications S/S for TT&C purposes, thespacecraft carries:

– an X-band high data rate transmission system to support high accuracyEarth-to-Moon ranging estimates—which is also a component of the Moongravitometric experiment—and the downloading of on-board stored B/Wand color images generated by two dedicated cameras for the geochemicalexperiments; To convey data over such a great distance, a 20-W TWTAand a mechanically repointable flat antenna with 30 cm sides was base-lined. The elevation over azimuth positioner, fixed to the lower edge ofthe antenna, allows to cover a 2π solid angle. This is necessary to estab-lish a link with Earth from any point of the Moon hemisphere visiblefrom Earth.

– an S-band low datarate system implementing a two-way intersatellite link(ISL) between the mother satellite and the released subsatellite. The ISLserves to remotely control the subsatellite-housed instruments (mainly theaccelerometer) and to implement, via high resolution and accuracy range-rate measurements, the gravitometric experiment. Since the subsatellitehas no attitude control devices it will move around the c.o.g. under theeffect of external torque-generating forces. In order not to perturb therange-rate measurements accuracy, positioning the antenna phase centeron the cube-like body center is highly preferred. Indeed, the S-band two-way transmission system puts some unconventional requirements to theantennas. On the mother satellite side a medium-directivity antenna is puton the bottom of the propulsion module looking in the direction opposite

Table 6 MAGIA spacecraft tentative mass budget

Major subsystem Mass, kg Remark

Propulsion S/S 33 DryStructure & Thermal 50.7 Truss-typeOBDH + PSE 29 Conventional technologyPower S/S 22.8 Incl panels, battery, harnessACS/S 13.2 RCS is part of Propulsion S/SComms 22.5 X- and S-bandPayloads 59.3 Incl P/L electronicsTotal dry 230.5Hydrazine 140 For orbit transfer mainlyN2 5 For corrections and AC S/STotal propellant 145Total wet mass 375.5

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to the velocity vector. The directivity is motivated by the need to reduce asfar as possible multiple reflections (multipath) from the Moon surface—which is being overflown at an altitude of only 70 km—when establishing aradiolink with the subsatellite kept at a relative distance of about 100 km.

10 Conclusions

The scientific MAGIA Mission can be implemented using a medium-smallsatellite platform which requires clever design expedients but does not implysignificant technology investments or risks.

The spacecraft mass budget is given in Table 6 and meets overall constraintstied to the self-imposed propulsion Subsystem mass allocation. A few topics,partly related to the payloads, partly to the Mission as a whole, and partlyto the spacecraft detail configuration and performance, need, nevertheless,refinements and deeper analyses.