C20014ExS

PROBA-3 Phase A Study Executive Summary Report

PROBA3-ASU-RPT-14Issue 01

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This document is produced under ESA contract, ESA export exemptions may therefore apply. These Technologies may require an export licence if exported from the EU

© Astrium Limited 2007

Astrium Limited owns the copyright of this document which is supplied in confidence and which shall not be used for any purpose other than that for which it is supplied and shall not in whole or in part be reproduced, copied, or

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PROBA3-ASU-RPT-14 Issue: 01 Page 2 of 20


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This Executive Summary has been compiled from the work of the PROBA-3 Phase A Study led by EADS Astrium. The members of the Study Team are listed below.

Astrium Ltd Andrew Davies Kelly Geelen Simon Grocott Stephen Kemble

Ronan Wall Carl Warren Alex Wishart

Astrium GmbH Klaus Ergenzinger

Astrium SAS Cyril Cavel Julien Morand

Verhaert Space Pieter Van den Braembussche Marline Claessens

GMV Lorenzo Tarabini

Swedish Space Corporation Nils Pokrupa

Laboratoire d’Astrophysique de Marseilles

Philippe Lamy Patrick Levacher

Sebastien Vives




CONTENTS

1 INTRODUCTION .....................................................................................................................................1-5

2 MISSION OBJECTIVES..........................................................................................................................2-6

3 TECHNOLOGY PAYLOADS .................................................................................................................3-7

4 SCIENCE PAYLOAD.............................................................................................................................4-10

5 MISSION ARCHITECTURE ................................................................................................................5-12

6 FORMATION FLYING .........................................................................................................................6-15

7 SPACECRAFT DESIGN ........................................................................................................................7-17

8 PROGRAMME DEVELOPMENT........................................................................................................8-19

9 CONCLUSION ........................................................................................................................................9-20




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1 INTRODUCTION This report provides an executive summary of the PROBA-3 Phase A study performed by a team led by EADS Astrium over the period July 2006 through July 2007. The Study was conducted in two parts. In part one ESA’s initial mission and system requirements were analysed and a baseline system design generated. This was reviewed with ESA at a Preliminary Concept Review held at ESTEC at the end of January 2007. In the second part of the study the baseline concept was studied in detail resulting in mission and spacecraft designs which were presented to ESA at the Preliminary Requirements Review held at ESTEC at the end of June 2007. A development plan and a ROM cost for the implementation phase has also been produced.

The Study team and the broad allocation of tasks are shown in Figure 1-1.

Figure 1-1: PROBA-3 Phase A Study team




2 MISSION OBJECTIVES The PROBA-3 mission is designed to provide in-orbit demonstration of new Formation Flying techniques and technologies being developed in Europe, validating their use for future operational Formation Flying missions. A series of Formation Flying manoeuvres will be performed which are designed to exercise the various technology items being flown. PROBA-3 will also carry a science payload in the form of an externally occulted solar coronagraph. This payload requires the two spacecraft to maintain a precise sun-pointing formation and this experiment will serve to validate mission in a quantifiable way. The technical demonstrations are summarised in Figure 2-1.

Formation Flying Techniques and Technologies Level of Demonstration

needed

FF Mission Requirement

Representation on PROBA-3

Position navigation to mm

Position navigation to micron

Position navigation to sub micron

High

High

As far as possible

Required

Variable

Variable

High

Medium

No

Absolute/relative attitude determination (arcsec)

Absolute attitude determination (sub arcsec)

High

High

Required

Variable

High

No

Vision based navigation High Variable No

Formation control High Required Medium

RF to optical metrology transition High Required High

Ground control deployment

Autonomous deployment

High

High

Required

Required

Medium

Medium

FDIR and anti-collision High Required High

Command and control sharing space/ground Strategies – master/slave, multi-master High Required No

RF metrology – coarse

RF metrology – fine

High

High

Required

Required

Medium

High

Optical metrology – coarse lateral

Optical metrology – fine lateral

Optical metrology fine longitudinal (DWI/FSI)

High

As far as possible

As far as possible

Required

Variable

Variable

High

Medium

High

Cold gas thrusters

EP (micro-ion)

High

High

Required

Required

No

High

ISL High Required High

Figure 2-1: Formation Flying technology on PROBA-3 – relevance to future missions




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PROBa-3 will also serve to develop the industrial and engineering processes for future formation flying missions. These are summarised in Figure 2-2.

Formation Flying engineering techniques to be demonstrated

Level of demonstration

needed

Required on FF Missions

Representation on PROBA-3

programme

Mission architecture design High required High

System design High required High

System modelling and simulation High required High

Formation Flying Test benches High required High

GNC test bench High required High

End-end validation High required High

Figure 2-2: Engineering process developments in PROBA-3

3 TECHNOLOGY PAYLOADS

3.1 METROLOGY

The Formation Flying metrology consists of 3 major pieces of equipment: RF Metrology system, an optical Coarse Lateral Sensor (CLS) and a High Precision Optical Metrology (HPOM) system which features a Dual Wavelength Interferometer (DWI) and a Fine Lateral Sensor (FLS). In addition to this each spacecraft uses star trackers to provide accurate attitude knowledge; this is particularly important because there is a very high degree of coupling between the attitude of the spacecraft and the measurement of the lateral separation of the spacecraft.

The RF Metrology system has been developed by Thales Alenia Space. This sensor is based on the use of GPS-like signals, although with different processing as both ends of the system are active. The RF metrology can operate in a coarse mode based on pseudorange measurements or in a fine mode based on carrier phase measurements. Coarse mode measurements are made with a single Rx/Tx antenna on each spacecraft. The range between the spacecraft is coarsely determined, and the angular position of the one spacecraft with respect to the other is very coarsely determined from the strength of the signal received. The Fine mode operates by using a triplet of antennas, 1 Rx/Tx and 2 Rx.

To provide 4π steradian coverage, these antennas are arranged in a tetrahedral pattern with four sets of antennas at the vertices. It is only strictly necessary to have fine measurement on the Coronagraph spacecraft when the Coronagraph and Occulter spacecraft face each other in the nominal formation configuration. Therefore, the Coronagraph spacecraft uses a triplet antenna at one vertex of the tetrahedron. These antennas can be seen in Figure 3-1. The RF metrology is primarily designed for missions outside the GPS sphere (for example at a Lagrange point). The maximum range is of order 30 km.




Figure 3-1: RF antenna configuration In fine mode the system should provide range accuracy to the order of 10mm and bearing accuracy (azimuth and elevation) of order 1o.

The optical CLS operates essentially as a very fine, narrow field of view star tracker, with an artificial star pattern produced on the Occulter spacecraft (the target) using a set of laser diodes. The CLS Optical head is mounted on the Coronagraph spacecraft (the chaser). The Field of View is wide enough to encompass the accuracy of the RF metrology, to allow transition between the two sensors. The lateral accuracy will be of order 1 arcsec.

Fine longitudinal and lateral metrology is provided by the DWI. The DWI is a heterodyne interferometer with a synthetic wavelength of 100mm. This permits the DWI to measure displacement on the order of 100 µm with an ambiguity of 100 mm. In addition, the laser signal of the DWI also impinges on a position sensitive detector (the FLS) via a beam splitter arrangement in the target retro-reflector which gives a precise measurement of lateral displacements. The DWI retro reflector is mounted near the centre of the Occulter disk and the optical head of the DWI is mounted near the CLS on the Coronagraph spacecraft. The HPOM is being developed by Astrium.




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Laser stabilisation

Laser 1 (f)

Phase Locked Loop

AOM1

AOM2

Laser 2 (f+3 GHz)

AOM4

AOM4

AOMs RF frequency generators

f + 80 MHz

f + 3083 MHz

f + 81 MHz

f + 3080 MHzLaser stabilisation

Laser 1 (f)

Phase Locked Loop

AOM1

AOM2

Laser 2 (f+3 GHz)

AOM4

AOM4


f + 80 MHz

f + 3083 MHz

f + 81 MHz


Laser 1 (f)

Phase Locked Loop

AOM1

AOM2

Laser 2 (f+3 GHz)

AOM4

AOM4


f + 80 MHz

f + 3083 MHz

f + 81 MHz


Laser 1 (f)

Phase Locked Loop

AOM1

AOM2

Laser 2 (f+3 GHz)

AOM4

AOM4


f + 80 MHz

f + 3083 MHz

f + 81 MHz

f + 3080 MHz

Figure 3-2: HPOM DWI working principle and breadboard unit

The laser wavelengths in both metrology systems are outside the forbidden range of 530-640 nm dictated by the coronagraph.

3.2 PROPULSION

The formation separation is controlled using electric propulsion micro thrusters with a dynamic range of order 100µN thrust. There are currently two types of electric propulsion thrusters under development by ESA which could be demonstrated on PROBA-3. These are the RIT-2/4 Radio Frequency Ion Thruster (RIT) being developed by EADS ST and the Miniature Gridded Ion Thruster (MiGIT) being developed by QinetiQ.

The Coronagraph spacecraft mass and power budgets for the Electric Propulsion option which are presented in §7 of this report were derived using the RIT2/4 thruster parameters. This is a more mature design than the MiGIT which is currently at a very low TRL.

It should be noted however that the intrinsically high specific impulse advantage of electric propulsion is lost in PROBA-3. To generate thrust on demand, the ionisation process for the thrusters must be continuously running, because the time from power on to ion production is on the order of 5-20 minutes. It would therefore be impossible to perform demanding control if the thrusters were powered down completely between every thrust application. Further, the required rapid variations in thrust level cannot be obtained through control of the propellant flow, and must instead be obtained by modulating the grid voltage. Control of the grid voltage essentially controls the exit velocity of the ions, and therefore affects the specific impulse of the thruster. This leads to the situation where the propellant supply limits the mission lifetime, which negates the object of using electric propulsion.

The Occulter spacecraft is passive, using no propulsion during Formation Flying operations while the Coronagraph spacecraft is the active spacecraft using the electric microthruster. Higher thrust is desired to provide small Delta-V manoeuvres to provide safe perigee passage. This is achieved using a cold gas propulsion system using micro thrusters on the Occulter spacecraft. Both spacecraft use reaction wheels as attitude control actuators, with the respective micro thrusters providing torque to prevent saturation of the reaction wheels.




4 SCIENCE PAYLOAD PROBA-3 will carry a solar coronagraph science payload. This is in the form of a coronagraph instrument on one spacecraft and an external occulting disk on the other spacecraft. In operation, the two spacecraft fly in a precisely controlled sun pointing formation, at a nominal Inter Satellite Distance (ISD) of 150m, as shown in Figure 4-1.

Figure 4-1: Coronagraph Formation Flying Concept

The coronagraph will achieve unprecedented spatial resolution of the solar corona in to 1.075 solar radii.

Solar Corona spectral information:Fe XIV: 530.3 nm (coronal matter, 1.8x106 K)Fe X: 637.4 nm (coronal holes, 1.0x106 K)He I: 587.6 nm (cold matter, 1.0x105 K)

Coronagraph instrumentSolar Corona spectral information:Fe XIV: 530.3 nm (coronal matter, 1.8x106 K)Fe X: 637.4 nm (coronal holes, 1.0x106 K)He I: 587.6 nm (cold matter, 1.0x105 K)

Coronagraph instrument

Figure 4-2: Solar corona image and coronagraph optical box with electronics




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The successful operation of the solar coronagraph will serve to validate the Formation Flying demonstration in a quantifiable way. The formation flying requirements are driven by need to maintain the nominal occultation of Sun. At the nominal inter-satellite distance (ISD) of 150m these are:

• Lateral Position Error ±3 mm (equivalent to 4 arcsec attitude error)

• Longitudinal Position Error ±100 mm

The lateral accuracy, in particular, is difficult to achieve because the 4arcsec value includes an absolute term from error in the Coronagraph to Sun Line of Sight (LOS - the angle β in Figure 4-3) as well as a relative term (the angle θ in Figure 4-3) which is measured using the relative metrology (the optical CLS). The LOS is determined using a standard star tracker and ephemeris data which would give at best an achievable positioning accuracy of 5 arcsec. The solution proposed in the Phase A is to use a Shadow Position Sensor (SPS), which is part of the coronagraph instrument, in a calibration exercise performed periodically during the mission to remove the bias between the Star Tracker (STR) and the CLS.

Figure 4-3: Effect of absolute LoS error on measurement of lateral offset D The Formation alignment bias cannot be solved by calibration alone, however, because there are two optical lines of sight, one for the coronagraph optics and another for the HPOM. To operate these two instruments simultaneously would require them to be co aligned with a very high accuracy due to their limited FoV. This problem arises primarily due to uncertainty in the internal alignment of the coronagraph optics. The HPOM could in principle be fitted with a steerable mirror but this is not proposed because it would complicate the instrument design. In fact the coronagraph Formation Flying operation does not require the HPOM measurement of longitudinal range – the fine mode RF metrology gives sufficient accuracy (see § 6 below). Therefore there is no need to operate the HPOM and the coronagraph optics unit simultaneously and the Coronagraph Spacecraft power budget is sized accordingly.




5 MISSION ARCHITECTURE The elements of the mission architecture are shown in Figure 5-1.

Launch Segment

VEGARockot

Propulsion Module

CoronagraphTelescope

CEB

CCB

CCD radiator

M1

M3

FPA

M2

CEB

CCB

LEOP NetworkOcculter Spacecraft Coronagraph Spacecraft

Formation FlyingTechnologies

S-Band Ground StationRedu, Belgium

Launch Composite

LEOPApogee Raising Operational Orbit

Ground Segment

Flight OperationsControl Centre Payload

DataSegmentGround

CommunicationsNetwork

Formation FlyingDemonstrations

Electric Propulsion

Figure 5-1: PROBA-3 mission architecture

A number of selection criteria for the mission design have been analysed, based on spacecraft related issues and demonstration requirements. Primary spacecraft constraints are the fuel required to reach orbit and ground contact availability which in turn drives communications requirements.

Demonstration related criteria are:

• Time spent per orbit in low gravity gradient environment

• Time available to complete demonstration manoeuvres (formation slews)

• Delta-V for formation manoeuvres (i.e. between different formation modes)

• Delta-V for Non-Keplerian motion phases.

• Time in an orbit to complete a manoeuvre sequence (timeline complexity)

A Highly Elliptical Orbit (HEO) with a period of 72 hours has been selected as the optimum orbit for the PROBA-3 mission objectives.

The mission scenario assumes a VEGA launch into a low elliptical orbit (200x1100 km). The Lisa Path Finder Propulsion Module (LPF PRM) is then used to raise the apogee to around 160000 km altitude in a series of about ten orbit raising manoeuvres. A final PRM manoeuvre is performed at the target apogee to raise the perigee altitude to around 800 km. The nominal orbit parameters are summarised in Figure 5-2.




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Parameter Value

Nodal Period 3 days

Apogee Radius 168301 km

Perigee Radius 7178 km

Inclination 5°

Argument of Perigee Depends on launch date

Right Ascension -10° (TBC)

Figure 5-2 Nominal Operational Orbit Parameters

This gives more than 48 hours during the apogee pass at which the gravity gradient force between the two spacecraft is less than 100 microN. Most of the science operations will be close to apogee and good communications over this long apogee period are advantageous. The nominal ground station selected is Redu in Belgium and so a Northern latitude choice for the apogee gives good link opportunities.

Figure 5-3 shows the ground track of the initial operational orbit.

Figure 5-3: Ground track for first orbit




Figure 5-4

Orbit track (red) and equatorial projection (blue) for first orbit

Figure 5-5

Orbit trace (red) and equatorial projection (blue) over two year mission

Figure 5.5 shows the initial orbit and Figure 5-5 shows how the orbit evolves over two years under the influence of lunar-solar perturbations and the earth’s J2 harmonic. The squares are 1 earth radius edge-edge. At the end of the two years the inclination has risen from its initial value of 5o to over 30o.

Figure 5-6 shows the timeline of the operational orbit.

Figure 5-6: PROBA-3 operational orbit timeline




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6 FORMATION FLYING The system architecture is shown in Figure 6-1.

Figure 6-1: PROBA-3 Formation Flying system architecture

For each spacecraft, there are two main functional blocks, Formation Flying Management (FFM) and Mission Vehicle Management (MVM). FFM is defined as the aspects of control that require knowledge of the state of the formation, not simply one of the spacecraft. MVM consists of all of those components that are within the control of a single spacecraft. Nominally, the FFM resides on the Coronagraph spacecraft while MVM is resident on both spacecraft. However, there are aspects of the FFM that are required on the Occulter spacecraft as well as on the Coronagraph spacecraft. For the PROBA 3 mission, this is primarily required to satisfy the failure case of losing communication between the two spacecraft. If loss of communication occurs, it is necessary for the Occulter spacecraft to know what the state of the formation is in order to correctly perform collision avoidance.

Therefore, information must be passed over the Intersatellite Link (ISL) for the Occulter to possess the knowledge of the state of the formation. In a larger satellite formation this function could be more elaborate and permit continued operation of the mission in the absence of a failed spacecraft; however these aspects will not be addressed in the PROBA-3 mission.

Within the MVM on each spacecraft, there is interaction between the Attitude Control System (ACS), Failure Detection, Isolation and Recovery (FDIR). Collision Avoidance is a Formation level function.

There are five operational Formation Flying (FF) modes involving the Coronagraph spacecraft (CS) and the Occulter spacecraft (OS). These are summarised in Figure 6-2.




Mode Demonstration Sensors Propulsion Formation GNC

Acquisition Sets up for transition to fine mode of the FF RFM. The OS flies inertially

FF RFM in Coarse and Fine modes

Electric Propulsion on CS

Closed loop, non-Keplerian motion

Coarse Sets up for transition to the optical CLS. The OS flies inertially

FF RFM in Fine mode



Transition coronagraph, Darwin-type Formation slew, rotation, resize, roll. OS flies inertially.

FF RFM in Fine mode, optical CLS



Fine Darwin-type Formation slew, rotation, resize, roll. OS flies inertially

HPOM DWI and optical FLS



Perigee OS manoeuvres prior to perigee to set up cartwheel, leader follower etc evolution during LEO passage. CS flies inertially in this mode

FF RFM in Coarse mode

Cold gas on OS

Open loop, Keplerian. R-GPS used on ground to verify FF RFM measurements

Figure 6-2: Formation Flying modes for PROBA-3

The expected Formation Flying positional control performance in Transition mode (which includes the coronagraph) is summarised in Figure 6-3.

Longitudinal Accuracy (mm) Requirement

Long Term (post calibration sensor position tolerances and multipath) 10.22

Medium Term (thermo-elastic positional errors) 0.2

Short Term (sensor noise and spacecraft positional control) 6.32

Total budget 16.75 74

Lateral Accuracy at ISD = 150m (mm) Requirement

Long Term (post calibration STR pointing bias, ephemeris errors) 0.83

Medium Term (thermo -elastic deformations involving STR and CLS) 1.3

Short Term (mainly pointing, position control and CLS noise) 0.97

Total budget 3.09 3.15

Figure 6-3: Formation Flying position control performance in Transition mode

Closed loop non-Keplerian demonstrations will be performed in Fine mode (i.e. using all optical sensors) over a range of ISD from 25m out to 250m, the range limit for the optical sensors. The longitudinal control precision in Fine mode is expected to improve to of order 2mm. However the lateral precision will be about the same as in Transition mode, because although the FLS is more precise than the CLS its signal will be very noisy due to attitude jitter of the Coronagraph spacecraft.




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7 SPACECRAFT DESIGN Figure 7-1 shows the two spacecraft in the coronagraph Formation Flying formation and Figure 7-2 list their mass budgets.

Figure 7-1: Occulter SC and Coronagraph SC in coronagraph Formation Flying configuration

CORONAGRAPH SPACECRAFT Current Mass (kg) Contingency (kg) Maximum Mass (kg)Data Handling 18.6 1.9 20.5Power Subsystem 20.3 1.4 21.7Communications 8.3 0.7 9.0AOCS 9.9 0.1 10.0Stucture 138.9 23.6 162.5Thermal Subsystem 10.0 2.0 12.0Harness 8.00 0.4 8.40Mechanisms 4.00 0.6 4.58Propulsion 39.4 7.7 47.1PLATFORM TOTAL 257.4 38.2 295.6Payload 37.7 5.6 43.3Formation Flying Metrology 32.3 3.8 36.1PAYLOAD TOTAL 70.0 9.4 79.4DRY TOTAL 327.4 375.0System Margin 20 % 75.0DRY TOTAL (INCL. MARGIN) 450.0Propellant 2.5CORONAGRAPH SPACECRAFT MASS AT LAUNCH 452.6 OCCULTER SPACECRAFT Current Mass (kg) Contingency (kg) Maximum Mass (kg)Data Handling 18.6 1.9 20.5Power Subsystem 18.9 1.3 20.1Communications 8.3 0.4 8.7AOCS 9.7 0.1 9.7Stucture 40.4 6.8 47.2Thermal 10.00 2.0 12.00Harness 8.00 0.4 8.40Propulsion 11.5 1.0 12.4PLATFORM TOTAL 125.3 13.8 139.1Payload 8.0 1.5 9.5Formation Flying Sensors 20.7 1.8 22.5PAYLOAD TOTAL 28.7 3.3 32.0DRY TOTAL 154.1 171.2System Margin 20 % 34.2DRY TOTAL (INCL. MARGIN) 205.4Propellant 2.1OCCULTER SPACECRAFT MASS AT LAUNCH 207.5

Figure 7-2: Spacecraft mass budgets




The CS structural mass is too high a percentage of the platform mass. This is being driven by the PRM requirement that the two-spacecraft stack combination should have a first lateral frequency > 50 Hz at launch, in a configuration with a fixed base at the PRM-CS interface ring. There are several approaches which could be taken to reduce the CS structural mass, including removal of the top and bottom floors from the load path, redesign of the main load bearing structure, or possibly switching to a carbon fibre structure. Alternatively it might be possible to relax the lateral frequency requirement but this would require a coupled analysis of the PRM plus spacecraft stack.

The spacecraft subsystems are listed in Figure 7-3.

Subsystem Coronagraph SC Occulter SC

Structure Outer panels (Al face sheets) Inner panels (Al face sheets) Optical bench + Titanium feet LPF interface ring Separation ring (SAAB) Deployable panel, CRFP face sheets

Inner structure (Al face sheets) Outer structure (Al on sides, CFRP on top, back) Occulting disk (CFRP) Separation ring (SAAB)

Thermal MLI Coatings

Radiator Heatpipe Coatings

Mechanisms Hinges (Dutchspace) Solar Array Drive (Oerlikon) Hold down and release mechanism (frangibolt NEA)

Communications 2 x S-Band antennas (STT) MGA (RYMSA) TMTC electronics (STT)

S-Band antennas (STT) TMTC electronics (STT)

Data Handling System (DHS) ADPMS (Verhaert Space) ADPMS (Verhaert Space)

AOCS 4 x star trackers (DTU) 5 x MEMS Gyros (Systron Donner) Reaction wheels (Dynacon)

2 x star trackers (DTU) Gyros (Systron Donner) Reaction wheels (Dynacon)

Propulsion System Electric Thrusters (RIT-2 EADS ST) or Cold gas (Marotta thrusters, ATK tank)

Cold gas (Marotta thrusters,ATK tank)

Formation Flying Equipments

S-band antenna triplet (SAAB helix antenna) 3 x RX/TX S-band antennas (STT) 2 x RF electronics (Thales) GPS antenna + receiver (Astrium) CLS optical head + electronics unit (EADS Sodern) HPOM DWI (Astrium)

S-band antenna triplet (SAAB helix antenna) 3 x RX/TX S-band antennas (STT) 2 x RF electronics (Thales) GPS antenna + receiver (Astrium) CLS OEU + target (laser diodes) DWI target beam splitter + FLS

Power Battery (SAFT) PCDU (Verhaert Space)

Battery (SAFT) PCDU (Verhaert Space)

Solar Array Solar Cells (Galileo Avionica)

Instrument Coronagraph optics, electronics (LAM) Shadow Position Sensor (LAM)

Occulter Position sensor (LAM) ARaSS (LAM)

Figure 7-3: Coronagraph and Occulter spacecraft subsystems




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8 PROGRAMME DEVELOPMENT PROBA-3 will be a more complex programme both technically and industrially than previous PROBA missions. This will provide an opportunity to test the new industrial and engineering processes which will be necessary for future large operational Formation Flying programmes.

A programme organisation structure for the development and operations phases of PROBA-3 is shown in Figure 8-1.

Figure 8-1: Programme Organisation Diagram

The space segment is organised under an industry prime who is responsible to ESA. PROBA-3 is a programme in which all aspects of the space segment specification and design are closely inter-related, yet these activities are distributed among several members of the industrial team.

ESA would be responsible for:

setting mission and system requirements managing the ESA funded Formation Flying technology programmes to TRL 5 interface between PROBA-3 and nationally funded Formation Flying technology programmes interface between PROBA-3 and the coronagraph science payload (assumed to be nationally

funded) launch segment comprising VEGA and the PRM LEOP operations




The Industrial team would be responsible for:

mission and system design spacecraft specification, design, and AIT payload Interface Engineering and payload AIT management of the ESA Formation Flying technology developments from TRL 5 to TRL 6 system AIT specific ground segment facilities (e.g. mission simulator) operations support

The outline of the Phase B/C/D programme schedule is shown in Figure 8-2.

Slice 1 System Review/Event Date

PRR June 2007

Phase B KO January 2008

SRR April 2008

PDR January 2009

FF Technologies at TRL 5

Slice 2 System Review/Event Date

CDR July 2010

Formation Flying Technologies at TRL 6

FAR October 2011

Launch June 2012

Operations start September 2012

Figure 8-2: PROBA-3 schedule milestones

9 CONCLUSION The Phase A Study has shown that the PROBA-3 mission is technically feasible and a closed system design has been achieved which meets the mission and system requirements. The driving requirements are:

• The Coronagraph Formation Flying position control accuracies, in particular the lateral accuracy.

• The requirement on the Lisa Pathfinder PRM for the Coronagraph/Occulter stack to have natural frequency > 50 Hz.

• The requirement to have nothing protruding from the Occulter disk.

The Study has identified a number of areas for follow-on work which would retire technical and programme risks prior to the start of the implementation phase.


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