57th International Astronautical Congress 2006IAC06-A3.5.5

LEARNING TO DEFLECT NEAR EARTH OBJECTS: INDUSTRIAL DESIGN OFTHE DON QUIJOTE MISSION

Ian Carnelli1, Andres Galvez2, Franco Ongaro3

1 ESA-ESTEC, Advanced Concepts TeamKeplerlaan 1, 2201 Noordwijk, The Netherlands

Ian.Carnelli@esa.int2 ESA-HQ, Advanced Concepts and Studies Office

3 ESA-HQ, Technology Planning and Monitoring Department8-10 rue Mario Nikkis 75738 Paris, France

Andres.Galvez@esa.int, Franco.Ongaro@esa.int

Abstract

Near-Earth Objects or NEOs include both objects having a likely asteroidal origin, and extinct cometsorbiting the Sun in the near Earth Space, crossing the region of the inner planets. Because of their closeapproach to the Earth, NEOs are the population of the smallest Solar System bodies that can be accessibleto detailed physical investigations, but in the same time they represent also a potential threat to our planet.Although impacts of large objects with catastrophic consequences are extremely infrequent, size of few tensor hundreds of meters in diameter can cause severe damage. A direct ground impact is not the sole threatsince NEOs might be the origin to a large scale Tsunami whose consequences can exceed those of the IndianOcean in 2004. The Don Quijote mission has been proposed by the European Space Agency as an asteroid-deflecting experiment with both a scientific and a practical perspective in the context of the managementof the NEOs impact hazard. The primary objective of the DQ mission is to impact a given NEO with aspacecraft (Impactor) and to measure the resulting variations of the orbital parameters and of the rotationstates by means of a second spacecraft (Orbiter) previously operating in the proximity of the asteroid. Aradio science instrument carried by the Orbiter will be used for the precise measurement of the asteroid orbitand of its gravity field. The Orbiter will also perform measurements to determine the asteroid mass, size andsurface properties. Secondary mission goals have also been defined, which would involve the deployment ofan autonomous surface package and several other experiments and measurements. Three industrial teamshave been awarded a contract by the European Space Agency to carry out phase-A studies in preparationfor the detail design and development phases. This paper presents the main intermediate results of thisdesign activity.

Background

It is common knowledge that asteroids have beenimpacting the Earth since its formation, over fourbillion years ago. There is evidence, like the meteorcrater in Arizona or images of the Tunguska forestin Siberia, which reminds us the devastating theeffects such impacts can generate. Although theprobability of these events is very low (in the orderof one every two centuries for a 30-50 m size aster-oids [1]), there is unfortunately very limited prac-tical knowledge of the Near-Earth Object (NEO)threat mitigation. In particular, many conceptualstudies have been carried out proposing different(and sometimes exotic) mitigation strategies but nospecific technological plan have been put in place.

Nevertheless, the magnitude of the problem hasbeen recognized in many occasions.

The resolution 1080 of the Council of Europe hasprovided recommendations on this issue [2] and atask force on the subject was established in theUK [3]. The importance of international initia-tives to further our understanding on NEOs hasalso been highlighted in other international forumsat the highest level such as the UN COPUOS [4]and the OECD Global Science Forum [5]. As aconsequence, in 2000 the European Space Agency’s(ESA) long-term space policy [6] identified NEOresearch as a task that should be actively pur-sued by the Agency. In July 2002 ESA decidedto found, through the General Studies Programme

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(GSP), preliminary studies of six space missionsthat could make significant contributions to ourknowledge of NEOs [7]. Three of these were obser-vatory missions (namely Earthguard-1, EUNEOSand NERO), and three were rendezvous missions(SIMONE, ISHTAR and Don Quijote). Followingthe completion of the six studies, ESA’s NEO Mis-sion Advisory Panel (NEOMAP) was established inJanuary 2004. Its role was to assess the studies andprovide ESA with recommendations for future ac-tivities. As a result the Don Quijote (DQ) missionconcept was selected and given highest priority [8].

Using these recommendations as the starting point,ESA conducted a first assessment in the context ofthe Concurrent Design Facility (CDF) [9] in De-cember 2004. This study was carried out by a mul-tidisciplinary team of spacecraft engineers and spe-cialists, also with the support from JAXA1. Theobjective was to assess the feasibility of several mis-sion scenarios based on the Don Quijote conceptwhile understanding their cost and technical riskimplications. This analysis was then used as thebasis for a second assessment [10] (carried out inJuly 2005) with the intention to define a referencelow-cost mission scenario and prepare for the in-dustrial phases-A studies by setting the system re-quirements and their priorities.

Three parallel industrial phase-A contracts have fi-nally been awarded in March 2006. These are in-volving Alcatel Alenia Space, EADS Astrium andQinetiQ as prime contractors, supported by severalsub-contracted European companies, research cen-ters and universities. These studies, which will endin the beginning of 2007, are assessing the designsand trade-offs performed by ESA and putting for-ward alternative design solutions for DQ.

1 Industrial phase-A objectives

The industrial studies have been structured in sucha way that the first step is to review ESA’s NEO2study [10]. In particular, the attention has beenfocused on the assessment of the mission require-ments and their implications on the mission archi-tecture. Special emphasis has been put on an earlyintegration of simple impact models into the trajec-tory analysis tool. This is fundamental for a correct

1Japan Aerospace eXploration Agency

understanding of the impact implications on a sys-tem level.

The second step focuses on the design of bothspacecraft, together with the equipment and op-erations needed for the characterization of the de-flection. The whole mission operations scenario, in-cluding the ground segment, will also be finalized.

The last step addresses the identification of devel-opment risks and critical technologies in order topave the way to later phases of the mission. Withthe intention of producing a “cost-aware” design, amaximum reuse of existing and flight-proven tech-nology was stressed. In particular, by maximizingcommonalities between the two spacecraft, enhanc-ing the use of autonomous operations, and baselinethe use of small-class launchers (such as Vega), thegoal is to combine technology demonstration for asmall low-cost interplanetary mission with the tech-nology readiness for an eventual NEO mitigation.

2 The Don Quijote mission

The objectives of the Don Quijote NEO mission arethe following [2]:

• primary objective: to impact a given NEA andto be able to determine the momentum trans-fer resulting from the impact, by measuringthe asteroid mass, size and bulk density andthe variation of both the asteroid’s center ofmass orbital parameters and its rotation state.

• Secondary objective: to carry out an Au-tonomous Surface Package Deployment Engi-neering eXperiment (ASP-DEX) and performmulti-spectral mapping of the asteroid. An op-tional extension of this secondary objective isthe characterization of the thermal and me-chanical properties of the asteroid surface.

As a result, two system options have been defined:

• Option 1: DQ “Light” mission that will ad-dress the primary mission objective only. Thesystem comprises an Impactor and an Orbiter,the latter carrying only the minimum payloadneeded to accomplish the mission primary ob-jective (i.e. to measure the linear momentumtransfer resulting from a hypervelocity impacton the target asteroid).

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• Option 2: “DQ+” mission, addressing boththe primary and the secondary objectives. Toachieve this goals, the Orbiter spacecraft willcarry an additional simple suite of engineer-ing and scientific payloads, including the ASP-DEX.

2.1 Mission scenario

The mission will contain the following elements (seefigure 1). In both system options, the Orbiter

Figure 1: Don Quijote mission scenario

spacecraft (called Sancho) is the vehicle that, fol-lowing launch 1 and early orbit phase 2 , performsan interplanetary cruise 5 by means of a solar elec-tric propulsion system (SEP). It autonomously ren-dezvous with the target asteroid 6 and is insertedinto orbit around it 7 . It measures its orbital pa-rameters, the mass, size, gravity field and shape,before 8 and after 11 impact in order to assess themomentum transfer. In addition, the Orbiter shalloperate as a backup data relay 10 for transfer-ring the collected Impactor Guidance, Navigationand Control (GNC) engineering data, and imagethe impact from a safe parking position. In paral-lel to attaining this primary objective the Orbiter,in the “DQ+” mission option, pursues scientific in-vestigations of the asteroid, addressing part of themission secondary goals. Finally, after completionof the primary mission 12 , the “DQ+” Orbiterwill carry out the ASP-DEX 13 and act as datarelay for the surface package 14 .

The Impactor (called Hidalgo) is the vehiclethat, after an interplanetary cruise with mini-mum ground segment (G/S) support 4 , will per-form completely autonomous terminal guidanceand navigation manoeuvres towards the target as-teroid ( 5 and 6 ). It relays GNC engineering dataand images of the target to the G/S 7 and Orbiter

spacecraft 8 , while impacting at very high relativespeed (∼ 10 km/s) against the asteroid’s surface.This spacecraft will demonstrate the autonomousGNC capability based on optical navigation.

As NASA’s Deep Impact mission already proved,vision-based autonomous guidance navigation andcontrol can be achieved. However DQ will be a pre-cursor mission as it will be facing additional chal-lenges linked to the target’s reduced dimensions.

2.2 Main mission requirements

The main mission requirements can be divided intofour groups considering operations, Impactor, Or-biter and technology.

From a system operations point of view, the twospacecraft must be launched separately using asmall class launcher (such as Vega) or medium classlaunchers only if the performances reveal to be in-sufficient. The Impactor is to be launched onlyafter the Orbiter successful rendezvous with the as-teroid and the ASP-DeX must be carried out onlyat the end-of-mission, not to compromise the mis-sion’s primary objective. Finally the Orbiter mustallow for at least two months of radio science oper-ations in order to precisely determine the asteroid’sorbital parameters before and after impact, henceassess the momentum transfer.

As far as the Impactor is concerned, without tak-ing into account any ejecta effect, the impact mustbe such that the asteroid’s semi-major axis varia-tion is greater than 100 m. Also it must acquirethe target two days before impact and perform au-tonomous navigation in order to impact within 50m (3σ) from the center of mass.

The Orbiter instead must be capable of measur-ing such deflection with a 10% accuracy and back-up data for the Impactor’s GNC. As a technologydemonstrator it shall prove advanced autonomy-capabilities during 30 days of interplanetary cruise,the rendezvous phase within 100 km distance fromthe asteroid, and at least four consecutive orbitswhen performing radio science measurements.

Technologically, all system elements must meet areadiness level (TRL) above 6 by mid 2007 exceptfor the Orbiter and Impactor’s components that areneeded for autonomous operations. The latter haveto meet a TRL ≥ 5 by mid 2008.

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Figure 2: Target asteroids heliocentric orbits

3 Target selection

The selection of an appropriate target for the in-ternal pre-phase A and industrial phase-A studieswas based on a set of NEO characteristics that aremost relevant for the Don Quijote mission design.These were defined by ESA’s NEOMAP and aresummarized in table 1. As a result of this analy-sis two targets have been pre-selected [11], asteroid2002AT4 (baseline) and asteroid 1989ML (backup),represented respectively in red and green in fig-ure 2. Relevant orbital and physical characteris-tics of 2002AT4 and 1989ML are summarized intable 2. 1989ML is more accessible than 2002AT4,

Orbit characteristics preferred rangeRendezvous ∆ V < 7 Km/sOrbit type AmorMOID large and increasingPhysical characteristics preferred rangeSize < 800 mDensity ∼ 1.3 g/cm3

Absolute magnitude 20.4− 19.6Shape nor irregularTaxonomic type C-typeRotation period < 20 hoursBinarity not binary

Table 1: Target selection criteria

2002AT4 1989MLOrbital period [yr] 2.549 1.463e 0.447 0.137i [deg] 1.5 4.4∆ V [km/s] 6.58 4.46Orbit type Amor AmorMOID large largeAbsolute magnitude 20.96 19.35Taxonomic type D-type E-type2

Diameter [m] 380 800Rotational period [h] 6 (assumed) 19

Table 2: Selected targets’ main properties

thus more favorable from an Orbiter design pointof view. However, perturbing its trajectory wouldbe more challenging due to its larger mass, there-fore the 2002AT4 scenario is the sizing case for theOrbiter while 1989ML is the sizing case for theImpactor. Adopting this approach based on twodifferent target bodies, it allows to have a robustdesign that can adapt to new scenarios (e.g. alter-native targets in later phases of the mission) andlarge uncertainties in the asteroid properties (e.g.in case a similar mission had to be used to asses areal threat from an unknown NEO).

4 Reference design

4.1 Orbiter transfer

This section describes the reference transfer to thisbaseline asteroid 2002AT4. The Earth escape willtake place in mid-March 2011 (see figure 3) withan escape velocity of 3.5 km/s. This will be pre-ceded by a typically lengthy escape sequence, whichmight take about 3 weeks. Arrival will occur inearly January 2015, almost 2.5 years prior to im-pact. Sancho, after a series of short drift-bys forinitial target mass estimation, will be inserted intoorbit around it. This will happen at the latest inmid-November 2016, when the asteroid is still atover 2 AU from the Sun. It is more than 6 monthsprior to impact, which in this scenario would takeplace on 1 June 2017. In total, the Orbiter will per-form three revolutions around the sun. Its missionwill last until around 6 months after impact in order

2based on the results presented in [12]

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Figure 3: Orbiter transfer trajectory to asteroid2002AT4

to complete the second Radio Science Experiment(RSE) and measure the change in the target’s semi-major axis. Hence, the total mission duration is 6years. The overall Orbiter mission is summarizedin Table 3.

The Solar Electric Propulsion system (SEP), asexplained later in section 5, is chosen from theSMART-1 spacecraft. It is a Stationary PlasmaThruster (SPT) with input power at 1 AU of 1.78kW. Due to the large heliocentric distance values,the SEP cannot be operated throughout the wholeorbit, two heliocentric revolutions are therefore re-quired in order to complete the transfer. These con-tain five thrust arcs, mostly around the perihelion.The total propellant consumption is 76 kg, whichis still within the SMART-1 tank capacity, allowingfor some margin. It is note worthy that the most

Launch vehicle DneprParking orbit 300 kmEarth escape date 4.3.2011V∞ 3.5 km/sms/c 450 kgArrival date 4.1.2015transfer duration 3.8 yearsXenon consumption 76 kgthrust time 6312 h

Table 3: Summary of Sancho’s transfer to 2002AT4

relevant thrust arcs take place at low heliocentricranges where the available power, specific impulse,and thrust levels are higher. Following arrival, theOrbiter spacecraft will remain in the vicinity of theasteroid.

4.2 Impactor transfer

Due to the orbital properties of the target aster-oid, a high-velocity impact does not require a Venusswing-by, as it is the case for the mission to 1989ML(which is not discussed here). The Impactor islaunched in late December 2015 that is after theOrbiter’s arrival. It performs one complete helio-centric revolution and, on the outbound arc of thesecond one, it hits the asteroid on 1 June 2017with a relative velocity of over 9 km/s (see fig-ure 4). Earlier launches would also enable possi-

Figure 4: Impactor transfer trajectory to asteroid2002AT4

ble transfers, starting in September 2015 and ar-riving as early as April 2017 at 13 km/s. However,a mission scenario in which the impact would takeplace around perihelion is favored due to the re-duced Sun and Earth ranges and the possibility toperform Earth-based observation campaigns of theevent. With an escape velocity of 2.26 km/s, nodeep space manoeuvres or swing-bys are required.Finally the total transfer duration is less than 18

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Launch vehicle DneprParking orbit 300 kmEarth escape date 20.12.2015V∞ 2.26 km/sms/c into escape 790 kgms/c without CPS 560 kgImpact date 1.6.2017Impact velocity 9 km/sTransfer duration 1.45 yearsDeep space maneuvers none

Table 4: Summary of Hidalgo’s transfer to 2002AT4

months, which is quite efficient, simple and fast.Table 4 summarizes the transfer properties, noticethat impact takes place at an Earth range of 1.64AU.

Thanks to the final approach on the outbound arcof the trajectory, the value for the sun-spacecraft-asteroid angle3 remains large throughout the finalapproach, facilitating autonomous navigation dur-ing the final approach. Thus, a major requirementfor the mission design is fulfilled.

4.3 Orbiter spacecraft

A re-use of the SMART-1 bus has been considered.Though this approach provides a good referencecase to assess mission costs reduction and the ma-turity of the technologies compatible with TRLs≥ 8, there are some limitations. These are mainlygiven by the availability of a single PPS-1350 en-gine, a fixed xenon tank capacity that limits thepropellant mass to 84 kg (at 50◦C)4, and finallythe given bus structure. In order to accomplish themission, the input power to the SEP requires anincreased solar array surface consisting in an extrapanel per wing. Also a completely different com-munication subsystem consisting of two-degree-of-freedom (DOF) steerable 70 cm high gain antenna(HGA), medium and low gain antennas and a UHFantenna for the communication with the Impactorduring targeting phase and the ASP are required.

Considering the extended payload set (i.e. theDQ+ version), the Orbiter’s system mass budget

3The angle describes the asteroid’s viewing conditions,for a value of 180◦ it appears fully illuminated.

4or 91 Kg of Xenon at 40◦C

Total dry mass 395 kgPayload mass 20.6 kgTotal propellant mass 96 kgTotal wet mass 491 kgInput power 1.7 kW

Table 5: Orbiter mass budget

is summarized in table 5. Sancho’s payload canbe considered to be quite basic. It is defined bythe navigation camera, the RSE and a radar al-timeter addressing the primary objective. DQ+ ishowever complemented by a set of scientific instru-ments dedicated to the secondary mission goal, seefigure 5. Also being part as this secondary objec-tive, the ASP-DEX will also enable to investigatethe mechanical properties of a small asteroid’s sur-face. This knowledge will be important in order todetermine the feasibility of coupling devices ontothe surface of an asteroid under microgravity con-ditions, which would be required for the implemen-tation of mitigation strategies relying on a directcontact with the asteroid. The ASP would thusbe part of the payload of the Orbiter, which wouldcarry and deliver it to the surface at the end of themission, from an orbit about the asteroid. This ap-proach has been taken to minimize the uncertain-ties related to the Orbiter operations during thedeployment of this payload.

4.4 Impactor spacecraft

The mission of the Impactor spacecraft is a pecu-liar one: the spacecraft should remain in a dormantstate during most of its lifetime until the last daysof asteroid approach where the autonomous guid-ance takes over and targets it toward the asteroid.

Figure 5: Orbiter’s payload

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Total dry mass 523 kgPayload mass 9 kgTotal propellant mass 1162 kgTotal wet mass 1694 kg

Table 6: Impactor mass budget

During the cruise phase only minimum functionsare required but before the impact all the sub-systems have to be up and functional with highlevel of reliability.

A major system design constraint is also on thespacecraft mass that (contrarily to what is nor-mally required) shall be above a certain thresholdto achieve the required asteroid orbit deflection andlower than the launch system escape performance.This implies that the DQ approach is applicableonly to relatively small target asteroids. Its func-tion is to perform the impact with the target as-teroid by means of autonomous GNC. Navigationfrom Earth will in fact be available only up to afew hours before the impact. The propulsion mod-ule is not jettisoned at escape but kept attachedduring the whole Impactor mission duration as bal-last. Clearly this strategy would impose specificconstraints in the GNC subsystem design. But itwould have the advantage of increasing the totalmomentum transferred to the target, thus maxi-mizing the chances to achieve the required 100 mvariation in the target semi-major axis.

The major design drivers for the Impactor are:the optical autonomous navigation system basedon advanced on-board computer and high resolu-tion camera, low-cost requirements able to matchTRL ≥ 6, and no moving appendages (solar arraysand antennas) to achieve stringent AOCS pointingaccuracies. The spacecraft mass budget is summa-rized in table 6.

No specific scientific objectives are assigned to theImpactor, hence only the camera to support au-tonomous navigation is considered as payload. Itis accommodated on one lateral panel of Hidalgo,together with the relevant electronics. A secondlateral panel, this time internal, is dedicated to theaccommodation of the COMM equipment. Figure6 shows the Impactor design both in the launchconfiguration inside the Dnepr fairing (on top ofthe propulsion module) and in the flight configura-

Figure 6: Hidalgo’s launch and flight configuration

tion. As it can be noticed, it is provided with body-mounted solar arrays and fixed high-gain antennato impart the required structural rigidity duringcritical targeting phase before impact.

5 Industrial design

5.1 Alcatel Alenia Space

The design under assessment by the Alcatel AleniaSpace (AAS) led consortium is based on the iden-tified main critical issues. These are (1) the radioscience experiment (RSE) that involves several el-ements of the Orbiter (e.g. radio subsystem, HGApointing mechanism, GNC subsystem . . . ) and thecalibration of the link delays between the space-craft and the ground station. Also, (2) the uncer-tainty in the masses of the largest asteroids (in themain asteroid belt) generates gravitational pertur-bations having frequency scales not very well sepa-rated from the orbital period of the target asteroids.This coupling is dangerous for a correct deflectionmeasurement, especially for the DQ mission whereseveral perturbation require careful modeling. Inthis respect the (3) gravitational acceleration of thetarget asteroid is fundamental and it requires a pre-cise knowledge of the spacecraft optical character-istics. Another challenge is represented by (4) theweak orbital stability when maneuvering in prox-imity of the asteroid, which requires a specific on-board autonomy. Finally (5) the Impactor’s termi-nal guidance before impact involves developing ad

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Figure 7: Alcatel Alenia Space design concepts

hoc control laws, elevated levels of autonomy, pre-cise spacecraft pointing, and enhanced processingresources.

Taking these elements into account, the two space-craft’s designs are being optimized with the objec-tive of minimizing the number of different modulesto be built (direct impact on cost) and maximizingthe re-use of subsystem and equipment between thetwo launches. Hence, two architectures are understudy (see figure 7): (1) the Orbiter on top of an in-dependent propulsion module, which is also used asthe propulsion module for the Impactor launch; and(2) the Orbiter on top of an integrated Impactorspacecraft that is used as the Orbiter’s chemicalpropulsion module.

In terms of modules number this gives for option(1): two propulsion modules (with possibly recur-rent design), one Orbiter module, and one Im-pactor module. For option (2) only two Impactormodule structures (with possibly many recurrentequipments) and one Orbiter module are insteadrequired. Figure 8 shows a preliminary Orbiterconfiguration above the chemical propulsion mod-ule for option (2).

Figure 8: AAS preliminary Orbiter design

Figure 9: EADS preliminary Orbiter design

5.2 EADS Astrium

In order to comply with the requirements of all mis-sion’s phases, namely the cruise by means of SEPsystem, rendezvous with the asteroid to perform acharacterization of the object (remote sensing andradio science), move to a safe parking position tomonitor the impact, return to a close orbit and fi-nally release the ASP, the Orbiter candidate con-figuration (i.e. under assessment) is a Dawn-likedesign (see figure 9). This involves moveable solararrays, a fixed HGA and instrument panel on oppo-site sides of the bus, and a SEP system and camerafacing opposite directions to avoid impingement.

As far as the Impactor design, the main driver isto attain a high momentum transfer while keep-ing the mission cost low. This translates in thechoice of a small launcher (e.g. Vega) coupled witha propulsion module to achieve Earth escape. Asa preliminary solution to maximize the mass andthus the momentum transfer the propulsion mod-ule is foreseen to remain attached to the Impactoralso after Earth escape. Here three different Im-pactor options are currently under study (see fig-ure 10) that are based on different propulsion mod-ule (PM) architectures: a so-called “Dead” PM, a“Zombie” PM, and an “Integrated” PM. In the fistconfiguration, the propulsion module is kept com-pletely passive in all mission phases after Earth es-cape. This “Dead” PM option has already been dis-carded, during the terminal approach it is in factnecessary to ensure the thrust authority throughthe spacecraft’s center of mass in all directions.

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Figure 10: EADS preliminary Impactor design

The “Zombie” PM involves limited changes to thePM, namely the thrust authority for terminal ap-proach is implemented by additional thrusters andthe thermal control is to be adapted to other mis-sion phases. This configuration has the advantagethat the PM can be procured independently andadapted for the specific purpose. Finally, the “In-tegrated” PM option is not only the most mass-effective design but it also offers the possibility tooptimize the spacecraft’s geometry to enhance thegeneration of ejecta, hence increase the the mo-mentum transfer. Also, it provides optimal solu-tions to several design challenges such as the pro-tection of navigation cameras against thruster im-pingement, the HGA placement and compatibilitywith a cost-efficient approach (based on a suitablePM heritage), and possible integration by the PMmanufacturer.

5.3 QinetiQ

The QuinetiQ-led consortium has also completedthe mission design driver identification phase. Theinter-relation of the mission phases between the twospacecraft as well as the RSE requirements uniqueto the DQ mission have been given highest prior-ity. This translates in requirements such as thespacecraft’s autonomy during interplanetary cruise,the platform’s stability during RSE operations (i.e.no parasitic ∆V), camera imagining rates (and im-plications on the communication system and datastorage), autonomous targeting system and plat-form stability (e.g. no camera “jitter”) of the im-pactor, and “burst” mode communication prior to

impact that drives the communication subsystemdesign (e.g. data rates, power . . . ).

As mentioned already in section 5.1, the RSE is thefundamental driver for the entire mission design.This defines the optimal times when the deflectionexperiment should be carried out. Some of the pa-rameters affecting its success are: a low Earth range(to ensure high signal/noise ration and thereforeobtain the the best measurement accuracy), a highSun range (to reduce solar torques on the Orbiter),and avoidance of Solar conjunction. These trans-late in optimal Orbiter arrival dates and thereforeImpactor’s departures, both defining the optimalmission trajectories.

As in the reference scenario (see section 4.3),SMART-1 is currently under evaluation as candi-date bus for the Orbiter spacecraft (figure 11). Thisapproach is suitable for implementation of a small,low-cost interplanetary mission. In fact, the sizeallows the use of a small/medium-class launcher, itis capable of carrying the DQ payload and incor-porates an SEP system capable of meeting all theorbital requirements. Some changes have alreadybeen identified, principally in the thermal, com-munication, GNC, and AOCS system. The GNCsub-system could largely benefit from the ESA’sPRISMA mission. By adopting the active orbitcontrol developed on a vision-based sensor, au-tonomous operations such as hovering, fly-aroundsand holding points in a rotating asteroid frame canbe enhanced.

Three concepts are being investigated for the Im-pactor: a platform based on the re-use of SMART-1, a bespoke solution and a re-use of a commu-nication satellite bus (e.g. GIOVE A). In all casesbody-mounted arrays, in order to provide adequate

Figure 11: QinetiQ preliminary Orbiter design

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Figure 12: QinetiQ preliminary Impactor design

stiffness for the terminal guidance, and the use ofsole chemical propulsion are foreseen. For the be-spoke design (see figure 12), a PROBA-like ap-proach is envisaged. This is based on the use ofa new structure coupled with space-qualified sub-systems to ensure low-cost design and develop-ment, centralized architecture around a powerfulon-board computer, and considerable flexibility to-wards mission specific configurations. The space-craft autonomy shall require only minimum groundoperations activities to ensure the nominal missionsuccess. Finally, a high degree of modularity is usedboth for the hardware and the software.

6 Conclusions

The Don Quijote mission combines a low-cost tech-nology demonstration for small interplanetary mis-sion with technology readiness for an eventual NEOmitigation. It therefore represents an excellent ex-ample of a “NEO precursor mission” that couldpave the way for an effective NEO deflection mis-sion, independently of the deflection strategy fi-nally being considered. Don Quijote will measurethe mechanical behavior of the asteroid as a whole,and determine the orbital deflection triggered bythe impact of the Hidalgo spacecraft at a very highrelative speed. It will also carry out measurementsthe asteroid mass and bulk density and constrainits mechanical properties. In addition to this, in-vestigations in the close proximity and the surfaceof a NEO would provide excellent opportunities forscientific research to be carried out.

The reference scenario demonstrates that the mis-sion can be accomplished within a low-cost tech-

nology demonstration program. Currently under-going industrial studies are putting forward inno-vative solutions in order to base the design on a“cost-aware” approach. These are based on a max-imum re-use of existing and flight-proven technol-ogy, identification of commonalities between theOrbiter and the Impactor, and on the developmentof high degrees of autonomy.

Finally, having independent mission elements, DonQuijote is a mission enabling a flexible implemen-tation strategy from the perspective of ESA. It of-fers in fact the advantage of distributing evenly thefunding effort as the development phase of Hidalgowould start only after that of Sancho is concluded.

Acknowledgements

The authors would like to thank Gino Bruno Am-ata, Andreas Rathke, Nigel Wells and their teamsfor their outstanding work and dedication to themission. The constructive support of Diego Esco-rial was also greatly appreciated. Special thanksto Alan Harris, Willy Benz, Alan Fitzsimmons, Si-mon Green, Patrick Michel and Giovanni Valsec-chi (NEOMAP members) for their enthusiasm andinvolvement in the Don Quijote mission. Specialthanks to all the members of the CDF study teamfor their work, especially the study team leader An-drea Santovincenzo, and Michael Khan and Paolode Pascale for their dedication to the project’s suc-cess.

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[10] ESA/CDF. A reference mission for diminish-ing threat from near-earth objects. CDF Re-port NEO2 CDF-39(A), ESA/ESTEC, August2005. Final Version.

[11] A. W. Harris, W. Benz, A. Fitzsimmons, S. F.Green, P. Michel, and G. P. Valsecchi. Targetselection for the don quijote mission. Technicalnote, NEOMAP, February 2005. Report toESA.

[12] M. Mueller, A. W. Harris, and A. Fitzsim-mons. Size, albedo, and taxonomy of potentialspacecraft target asteroid (10302) 1989ML.ICARUS, 2006. submitted.

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