Design of a micro-satellite for precise formation flying demonstration

Acta Astronautica 59 (2006) 862–872www.elsevier.com/locate/actaastro

Design of a micro-satellite for precise formation flyingdemonstration

Richard Sanchez∗, Patrice RenardEADS-Astrium, 31 av. des Cosmonautes 31402 Toulouse Cedex 4, France

Available online 30 September 2005

Abstract

This paper presents a joint CNES-EADS Astrium contribution to the ESA’s SMART-2 project.SMART-2, 2nd mission of the Satellite Missions for Advanced Research and Technology program, slated for launch in

2006, will test key technologies needed to develop two ambitious ESA missions:

• LISA (Laser Interferometry Space Antenna), an ESA cornerstone mission dedicated to the detection and observation ofgravitational waves; to be launched in 2011,

• DARWIN, another ESA cornerstone mission dedicated to the search of Earth-like planets; to be launched in 2015.

In Phase A study of this demonstrator, one of the options contemplated by ESA was considering two formation-flying satellites.In that sense, and in order to both reduce and share cost, CNES proposed with the technical support of EADS-Astrium, tobuild one of them from its Myriade micro-satellite product line, mainly used for LEO scientific applications.

The study carried out has permitted to validate the concept of using a low-cost micro-satellite in a scientific interplanetarymission requiring not more than 10 �m inter-satellite position accuracy!© 2005 Elsevier Ltd. All rights reserved.

1. Scope of the �Sat definition for SMART-2

The study was focused on the Darwin technologydemonstration part of the SMART-2 mission, with ob-jective to validate the capability of two satellites to flywith accurate location from each other. To achieve thisobjective, the proposed micro-satellite had to achieve

∗ Corresponding author.E-mail addresses: [email protected]

(R. Sanchez), [email protected] (P. Renard).

0094-5765/$ - see front matter © 2005 Elsevier Ltd. All rights reserved.doi:10.1016/j.actaastro.2005.07.039

three major functions:

• Radio-frequency metrology;• Optical metrology;• Guidance, navigation and control (GNC).

On one-hand the metrology was providing the GNCwith distance measurement within the required accu-racy. On the other hand, critical sensor/actuators tech-nologies were to be implemented in the GNC to con-trol such accuracy.

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R. Sanchez, P. Renard / Acta Astronautica 59 (2006) 862–872 863

Thus, the major challenges for the Myriade platformwere mainly related to

• the accommodation of the radio-frequency unitsand the optical bench of the high precision opticalmetrology (HPOM),

• the management of the coupling effect between theattitude control system and the metrology unit mea-surements (with end-to-end performances require-ment of 10 �m inter-satellite position accuracy),

• the use of electrical propulsion system (FEEP),• the configuration during orbit transfer (concept of

cargo and �Sat passenger).

2. Darwin mission requirements summary

This chapter describes the major Darwin phases.

2.1. Formation deployment

This phase covers the operations from separation ofthe eight Darwin spacecraft from the launcher(s) untila coarse constellation is created with the accuracy of10 m at inter-satellite distance > 250 m. At the end ofthis phase, the flyers and the master spacecraft havereached their nominal positions relative to the hub witha 1 cm3 uncertainty; a configuration allowing subse-quent laser metrology enabling and fringe acquisition.

The modes described hereafter use of RF sensingonly for constellation navigation.

2.2. Fringe acquisition

The first step after deployment consists instabilizing the baseline and improving the lateralpositioning by using laser metrology (ranging andcoarse lateral sensor); in order to enable inter-satellitebeam exchanges and to reach the inter-satellite dis-tance (ISD) stability necessary for fringe acquisition.The absolute pointing of the flyer spacecraft is alsoimproved by shifting from the star sensors to thetelescope signal.

At GNC level, the optical path delay (OPD) is dueto two independent errors (RSS combination):

• An uncertainty on the inter-satellite distances<1 cm;

• A flyer transverse offset relative to the interfer-ometer plane defined by the observation direction

Towards optical target

∆ OPD

Towards optical target

Fig. 1. Correct and tilted constellation configuration (2D example)creating external optical path delay (OPD).

<1 cm. Indeed this offset induces an OPD error(so-called “external OPD”) due to the tilt of thebaselines relative to the observation direction (tele-scopes line-of-sight) as shown in Fig. 1.

The uncertainty on OPD obtained from RF sensingand coarse transverse optical metrology is then 1.4 cm.

2.3. Nulling interferometer mission

During normal operation, the beams transmitted viathe telescope flyers must be co-phased for recombina-tion in the central hub. The OPD between the beamsfrom a pair of telescope flyers must be <20 nm, splitin 5 nm allocated to control error and 15 nm to instru-ment systematic errors.

The interfering beams must also have balanced( 1

1000 th) intensities, which requires to have beam su-perimposition on the fringe tracker with the sameaccuracy ( 1

1000 th) of the diffraction spot of 1.22 �/D,where D is the telescope diameter). This correspondsto 0.08 mas on the detector for 1.5 m telescopes.

2.4. Imaging interferometer mission

It is expected that OPD and pointing requirementswill be significantly relaxed for imaging (by at least a

864 R. Sanchez, P. Renard / Acta Astronautica 59 (2006) 862–872

factor of 10). Indeed, conventional OPD requirementsfor interferometric imaging are in the range of �/10to �/100, and beam superposition error only affectsfringe contrast. Consequently, an accuracy of 1

10 – 1100

of the diffraction spot is sufficient. As a starting point,1

100 th is assumed.The major difference compared to Nulling is that

the interferometer baselines shall be continuouslyvaried (length and orientation in the interferome-ter plane) during observations to achieve completescanning of the Fourier plane within a reasonableduration.

3. SMART-2 and Darwin compared requirements

The study considered that all technologies requiredfor Darwin would not be tested on SMART-2. Someof them would be verified by ground-based measure-ments. The flight tests would concentrate on forma-tion deployment, reconfiguration and precision forma-tion flying for the two scientific modes, Nulling andImaging.

Although the analysed SMART-2 option was com-posed of only two satellites, all the formation flyingkey technologies required for DARWIN, could, how-ever, be validated.

The foreseen sequence of modes was:Orbit transfer:

• Transfer phase from Earth up to L1 Lagrange point;

Formation deployment:

• Stabilization of attitude and on-station control• Initial formation acquisition and collision avoid-

ance quite similar to Darwin;• Coarse baseline control mode using star tracker and

RF subsystem measurements:range: 25–250 m; range accuracy: 1 cm; attitude ac-curacy/inertial frame: 0.5◦.

Nulling interferometer mission:

• Acquisition of the 1-cm position accuracy in 3D.With only two satellites this would require line-of-sight measurements with a refined accuracy from atransverse laser metrology system (e.g. a divergentlaser associated with a CCD detector).

Fringe acquisition:

• Length freezing mode based on laser metrology,down to the accuracy needed for fringe sensor ac-quisition.

• Fine control mode using the fringe sensor.

Imaging interferometer mission:

• Reorientation and resizing between Nulling missionobservations and during Imaging mission observa-tions.

Table 1 compares GNC requirements values betweenDarwin and SMART-2.

4. SMART-2 technology requirements

4.1. RF metrology

The RF metrology system is a major technologystep forward required for Darwin, the performancesof which cannot be assessed on ground because ofthe difficulty to reproduce the conditions of opera-tion in space (relative dynamics, RF environment). RFmetrology shall therefore be validated to the maxi-mum extent by the SMART-2 demonstration. The RFmetrology will be used in all Darwin-related modesand for the initial deployment.

The “basic hardware set” of the RF subsystemproposed in the study was a pair of “transceivers”(emitter/receiver) and a set of emit/receive antennasflown on the two satellites with the following twofunctions:

• 2-Way pseudo-ranging functions: each transceiveremits a GPS-like signal carrier modulated with abinary code (the navigation message). The oppo-site transceiver receives this signal, demodulatesthe binary code and measures its delay relative toan internally-generated code. When multiplied bythe speed velocity, this delay provides a pseudo-range measurement. Their respective pseudo-rangemeasurements are then exchanged between thetransceiver (this is autonomously achieved throughthe navigation signal) to resolve the clock biasbetween them. The ISD information is therefore


Table 1SMART-2 and Darwin compared GNC performance requirements

Mode Parameter Darwin SMART-2requirement requirement

Deployment mode Minimum inter-sat. distance 20 m 20 m(coarse optical Sun direction 40◦ half-cone from 45◦ half-cone fromsensors) solar array normal solar array normalBaseline Relative 25–250 m, 25–250 m,control mode distance selectable selectable(RF metrology Relative <1 cm & 1 mm/s 1 cm, az & el err.sensors) positioning 3D <0.25◦ (half cone)

Absolute <1 arcsec & Use off-the-shelfpointing 0.1 mm/s sun sensors formation roll <0.5◦

Fringe Scan duration 400 s 400 sacquisition Fringe sensor 0.1 Hz 1–10 Hz(optical sampling ratemetrology Scan step 4 �m at � = 1 �m 5–50 �mperfos) 50 �m at � = 4 �m

Scan rate 0.4 �m/s at 50 �m/s� = 1 �m5 �m/s at � = 4 �m

ISD absolute ±80 �m 0–2 cmmismatchISD stability 2.8 �m >5.50 �mISD drift 0.28 �m/s >50 �m

OPD stability 5 nm ±1 �mGlobal tilt 0.002 mas @ 250 m 1◦

(corresponding to 5 nm OPD)Nulling Absolute 8 mas 10 arcsecinterferometry attitude to the flyers(optical Flyer relative 0.1 mas/s Not applicablemetrology perfos) angular velocity

Relative hub-flyer attitude 60 mas 15 arcsecRelative hub-flyer 1 cm 3D 0–2 cm ISDpositions az & el: 0.25◦ half coneRelative hub-flyer velocity 1 mm/s 3D ISD : 1 mm/s

Imaging Baseline <3.75 mm/s <3.75 mm/sinterferometry change rate(optical Duration of TBD s TBD smetrology imaging periodsperfos) Resizing ISD : 25–250 m Same as Darwin

and return�ISD <1 m in 1 sDuration �16 h

available on both satellites at any time. This func-tion requires dedicated emission/reception RF an-tennas and relevant transceivers. In order to coverthe whole space, two antennas placed in opposite

locations need to be implemented on each satellite(Fig. 2);

• Transverse position (azimuth/elevation) deter-mination function: this function is in charge of


Z

Mas terDT P µSat

Flye rLTPMas ter

DT P µSatFlye r

LTPLi sa

Da rw in

µ Sa t

Fl ye rMas ter

DT P µSatFlye r

LTPterDT P µSat

Flye rLTPLisa

View from X

Transmit

Receive

Darwin µSat Flyer

Precision Flying Axis (Xsat)

Fig. 2. RF antenna implementation on SMART-2.

determining the direction of the other satellite(defined in S/C frame through elevation and az-imuth angles). The principle of so-called “RFgoniometry” is to measure the direction of the in-coming RF signal using the differential phase ofthe carrier signal received by three antennas (Fig.2). The accuracy is directly proportional to the dis-tance between antennas, and the main error sourceis multipath (reflection of the RF signal on nearbyobstacles), so the antennas need to be implementedclose to the edges of a satellite wall with minimumobstacles between them. When combined with themeasurements of an absolute attitude sensor (e.g. astar tracker) on both satellites, this function allowsre-building of the 3D relative motion of one space-craft relative to the other, in the reference inertialframe. Let us note that having a star tracker oneach satellite and RF goniometry on only one endis sufficient to determine the relative position. RFgoniometry on both ends would also provide therelative attitude between the S/C, of no interest forSMART-2. Therefore, the “minimum” RF metrol-ogy configuration proposed for SMART-2 wouldimplement the RF goniometry function only on theLISA satellite, to maximize possible antenna sepa-ration and reduce subsystem mass/power demandon the Darwin �satellite.

4.2. Optical metrology

SMART-2 had to demonstrate the feasibility of for-mation flying with accuracy in the nm range. Such

extreme performances call for precise and complexsystems. For this purpose, different optical equipmentswith optimised dynamical range and accuracy had tobe accommodated.

Two different metrology levels have to be con-sidered in a Darwin type experiment. The internalmetrology that freezes the constellation configuration,and the external metrology which ensures the correctpointing of the interferometer onto the guide star.SMART-2 experiment was representative of the firsttype of metrology.

The line-of-sight acquisition and pointing couldbe done with high accuracy by means of classicalequipments such as CCD camera and low powerlaser source. These equipments were covering the25–250 m distance range between the two satellites.The inter-satellite distance measurement accuracywas much more demanding and was driving directlythe choice of laser metrology technology. The threemain parameters sizing the laser were: the length ofcoherence of the laser source, the 1 �m accuracy andthe capability to generate different wavelengths forabsolute distance measurements. For SMART-2, alow coherence source associated with two fixed delaylines with lengths of 25 and 250 m to cover the ISDrange were selected (Fig. 3).

The �Sat metrology hardware was then able to pro-vide the following accuracies:

• Coarse lateral position of the two satellites, approx-imately 1 mm;

• Fine lateral and longitudinal distance between thetwo satellites, 10 �m;

• Fringe position in the interferometer defined by aninternal fiber reference and the counter satellite,1 �m.

It was also possible to develop a simple fringe-trackingsensor, but this approach was found to be less repre-sentative.

4.3. Micro-Newton propulsion system

The micro-thrusters used for drag-free and attitudecontrol during the LISA demonstration could also beused for accurate pointing and relative position ofthe two satellites of the Darwin demonstration. Asa consequence, the �N-thrusters on the �Sat had to


Co ar se La te ra lSe ns or

ite Li ghtsour ce

Le ng thRe fere nc e

Fr in geTr ac ke r

La ser So ur ce Lo ng itud in alSe ns or

(1 0 µm )

Co ar se La tera lSe ns or

Re fere nc e

Le ngt h

Fri ng e

Tr acker

ser S our ceLo ngi tudi na l

Se ns or(10 µm )

Co ar se La te ra lSe ns or

Le ng thRe fere nc e

Fr in geTr ac ke r

La ser So ur ce Lo ng itud in alSe ns or

(1 0 µm )

DARWIN µsat

UHF Link

Coarse Lateral Sensor Collimated

Laser

LISA mini-satellite

Longitudinal Sensor (10 µm)

Laser Source

Reference Length

FringeTracker

White Light source

Delay Line

FineLateralSensor(10 µm)

Fig. 3. Optical metrology implementation on SMART-2.

provide a 6-DOF control capability with tolerance tothe failure of at least one thruster (or one cluster if ar-ranged in clusters). The required authority was in therange of 100 �N.

Another challenge was the capability of the �Sat tosupport long consumption from FEEP and the use ofthis electrical propulsion as unique propulsion modulefor all modes, in particular on station setting after theorbit transfer.

The thrust noise was the driving factor as opposedto thrust vector misalignment, bias and scale factor.

In order to derive thruster noise requirements, a pre-liminary allocation of the ISD error contributors wasconsidered assuming equal quadratic weighting:

�ISD = �ISDsensors︸︷︷︸

1/√

2

+ �ISDsolar_pressure︸︷︷︸

1/2

+ �ISDthrusters︸︷︷︸

1/2︸︷︷︸

1/√

2

(1)

The thruster noise transmission in the GNC loop wasevaluated as a function of the control bandwidth.

(LISA)

Smart-2 specificationDarwin specificationSolar Pressureqthrusters=1

qthrusters=0.1qthrusters=0.01

10-3

102

100

10-2

10-4

10-6

10-8

10-2 10-1 100 101

Controller cut-off frequency (Hz)

ISD

err

or (

µm)

Fig. 4. Thruster noise impact on the inter-satellite distance (ISD).

Fig. 4 compares OPD requirements for SMART-2 andfor Darwin (corrected by the SMART-2 and Darwinmasses ratio, and assuming the same error allocationfor Darwin), and the OPD caused by solar pressure


Fig. 5. SMART-2 �Sat in stowed configuration.

and various thrust noise levels (specified in �N/√

Hz)as functions of the controller cut-off frequency. Themajor conclusions were:

• the rejection of the solar pressure can be achievedwith very low control bandwidth: <0.001 Hz (about0.05 Hz for Darwin);

Fig. 6. SMART-2 �Sat in orbit configuration.

• the dominant disturbance is the thruster noise con-sidering requirements of 0.1 �N/

√Hz, as for the

LISA demonstration. The resulting control band-width of about 0.01 Hz (about 0.5 Hz for Darwin)is significantly higher but still compatible with thethruster technology.

5. SMART-2 �Sat configuration

The mechanical architecture coped with the Myr-iade structure to enhance the re-use and associatedcost gain.

It is worth noticing that the design was defined with:

• two batteries;• two UHF transponders and relevant antennas

(optional);• two RF metrology units and relevant antennas (with

possibility to have them internally redunded to pro-vide additional mass margin). On the other hand,it was assumed that the same antenna was usedfor signal transmission and reception, and that onlyone antenna was required for the Darwin �satellite(baseline for pseudo-ranging function only); finally,two identical antennas were embedded to performomni coverage;

• one emergency transponder and one antenna(optional).


Fig. 7. SMART-2 �Sat internal layout.

The computed mass budget shows 144 kg associatedto a system margin of 14%.

Views of the SMART-2 micro-satellite are presentedin Figs. 5–7.

6. SMART-2 �Sat functional description

6.1. Overview

The main challenge was to deal with a �Sat con-cept together with highly challenging performances.The result led to implementing at �Sat level all func-tions relevant to scientific mission performances, andto relying on another satellite (minisat) for all missionservices (Earth link and transfer phase handling forinstance). This approach also offers new perspectivesfor other missions by using a system made of severallow-cost �Sats in orbit, controlled in position, associ-ated with a minisat (mother ship) with secured archi-tecture offering reliable links with Earth and extendedlifetime for mission follow-on.

The SMART-2 �Sat electrical architecture de-scribed in Fig. 8 was built on the Myriade avionicdesign. The Myriade redundancy philosophy is nom-inally to embed only redundant functions for theTM/TC communications and some key devices insidethe main on board computer (OBC). This philoso-phy was kept and remained in line with ESA rec-ommendations that did not require redundant units.However, some redundancies were proposed for somecritical functions as pseudo-ranging function andinter-communications for TM/TC link (Table 2).

6.2. Power system

Table 3 shows the power budget used for the sizingof the battery and the solar array. Using a recurrentSolar array from Myriade provided a margin of 17%.

6.3. Data management

The on board data-handling sub-system wasachieved around the central OBC. All the payloadunits were connected through the standard I/F (RS422and UART RS422) in order to exchange the telecom-mand and the telemetry data. Other equipments suchas FEEP, S band and UHF units were connected tothe RS422 I/F as well.

As all data exchanges were identified at lower rate(1 Hz), the centralized architecture around the OBCwas fully compatible with the Myriade data-handlingcapacity. Depending on the detailed real-time con-straints of each sub-system (house keeping and


Fig. 8. SMART-2 �Sat electrical architecture.

payload), the detailed allocation of each unit vs eachOBC I/F still had to confirm the real-time perfor-mances.

This architecture also could be simplified as theS band transponder was an option, and as the UHFtransponders could be removed if the RF payload unitwould be able to support the TM/TC link.

6.4. AOCS

The selected star tracker was a conventionalmedium field-of-view sensor providing autonomouslythree-axis attitude determination relative to the iner-tial reference with a better accuracy about the twoaxes perpendicular to the star tracker line-of-sight. Afactor of five geometrical amplification for the attitudeestimation error around the line-of-sight was assumed.

Star tracker of type SODERN SED 16, or ASCfrom the Denmark Technical University, were com-pliant for the best axes with the ESA specification(10 arcsec @ 3�).

The preliminary simulations of the SMART-2ACS/GNC system designed for the Darwin demon-stration allowed to achieve the following points:

• the complex transition between coarse RF metrol-ogy to fringe acquisition is feasible with the avail-able technology;

• fine formation-flying and OPD performance re-quirements are met with margins for the twoinvestigated fringe acquisition and OPD controlconcepts (with and without delay line);

• the proposed architecture with maximum decou-pling between mini and micro satellites is satisfac-tory;


Table 2SMART-2 �Sat redunded and non-redunded functions

Redunded function Justification Units

Inter S/C In order to prevent Redundant UHF RX/TXcomm’s collision and distance

limit exceed

RF Same justification as Redundant RF metrologymetrology per inter S/C coms unit

FEEP Same justification as Four FEEP embedded andper inter S/C coms system robust to the lost of one

Non-redunded function Justification

Emergency As this function should have been used onlycomm’s with following a major failure on the use of the interground S/C link, the need of a redundant transpondersegment was considered to cover double failure which

was largely out of scope

Power distribution Myriade design

Star tracker Myriade design

Star sensor Myriade design

OBC Myriade design

Table 3SMART-2 �Sat power budget

Equipment Qty LEOP Safe Deploy. Observ. Reorient. Includedmode margin (%)

OpticalMetrology 1 35.1 35.1 15RFMetrology 1 0.00 18.0 18.0 18.0 18.0 20DC/DCconverter 1 0.0 5.4 5.4 15.9 15.9 naTotalPayload 0.00 23.40 23.40 69.00 69.00Power PCDU 1 9.90 9.90 9.90 9.90 9.90 10AOCS Star tracker 1 0.00 0.00 8.40 8.40 8.40 5FEEP 0.00 48.00 48.00 48.00 48.00 20Command & On-Board processor 1 5.25 5.25 5.25 5.25 5control UHF emitter/receiver 1 3.60 3.60 3.60 20

S-band transmitter 1 6.00 20S-band receiver 1 3.36 20

Thermal Heaters 15.00 6.00 3.00 3.0 3.00Harness 0.80 0.80 0.80 1.7 0.90

Total 25.70 102.71 102.35 148.85 148.05


• the performance drivers have been identified: cou-pling of laser metrology sensors (longitudinal andlateral) with the pointing error, �N thruster noiseand response time, fine longitudinal metrology res-olution.

• the lateral formation-flying performance (as mea-sured by the fine lateral sensor) is hidden by theimpact of �Sat pointing stability error.

6.5. Propulsion

The SMART-2 requirements matched the �N FEEPthrust capacity for all satellite phases, authorizing thusa fully recurring FEEP system.

The FEEP thrusters were sized by reconfigurationmanoeuvres and orbit corrections rather than by exter-nal disturbances (and therefore not by satellite config-uration). It was demonstrated that a thrust of 100 �N,including significant margins (to account for mod-elling uncertainties and different manoeuvre profiles)was sufficient in particular with power sizing. 150 �Nwas, however, specified in the ESA phase A study, totake into account the needs for the LISA demonstra-tion, and was kept as baseline.

The evaluated Nitrogen mass was safely sized at5 kg for a 2-year mission. Due to that low value, coldgas thrusters were also seen as good candidates.

7. Conclusion

The study carried out by astrium on request ofCNES has permitted to rapidly design the mechan-ical, thermal, electrical architectures, and most ofall the avionics, of a micro-satellite meeting the

requirements of a Darwin key technologies demon-strator, in a SMART-2 option made of two formation-flying satellites.

It was clearly demonstrated that the use of �sats forinterplanetary missions is of interest, at the cost of aso-called “service satellite” in the range of a minisat,to support in-orbit transfer and TM/TC link with Earthstations.

While the SMART-2 second satellite is alreadyscheduled for LISA technology demonstration, thismicro-satellite based on the CNES Myriade productline presents a solution optimising the cost for theDarwin features to be demonstrated.

The main critical items: RF metrology, opticalmetrology and FEEP have been studied thoroughlyand their implementations in the �Sat validated.

Finally, this study has proven the feasibility of us-ing recurrent platforms for modular scientific appli-cations with very stringent requirements, such as aninter-satellite distance of 10 �m for instance. Such re-sult is encouraging the use of low-cost solutions evenin the case of high-valued missions.

Further reading

[1] L. Vaillon, J. Lebas, E. Sein, B. Calvel, Formation flying systemdesign for the Darwin demonstration of the Smart-2 mission,International Symposium on Formation Flying, Mission &Technologies, October 2002.

Design of a micro-satellite for precise formation flying demonstration

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