PROBA3 CDF Study Report: CDF-42(A) December 2005 page 1 of 231 CDF STUDY REPORT. Formation Flight Technology Demonstrator and Coronagraph Mission

Transcription

1 CDF STUDY REPORT Formation Flight Technology Demonstrator and Coronagraph Mission CDF-42(A)

2

3 page 1 of 231 CDF STUDY REPORT Formation Flight Technology Demonstrator and Coronagraph Mission

4 page 2 of 231 FRONT COVER satellites flying in formation for Sun-corona observation

5 page 3 of 231 STUDY TEAM This study was performed in the ESTEC Concurrent Design Facility (CDF) by the following interdisciplinary team: TEAM LEADER D. de Wilde, TEC-MCS COMMUNICATIONS P. Holsters, TEC-ETC POWER S. Zimmermann, TEC- EPS CONFIGURATION R. Westenberg, TEC- PROGRAMMATICS/ O. Brunner, TEC-TCC MCS AIV COST P.Ponzio, TEC-SYC PROPULSION J. Gonzalez del Amo, TEC-MPE D. M. Di Cara, TEC-MPE DATA HANDLING G. Furano, TEC-EDD RISK F. Restagno, TEC-QQD GNC F. Ankersen, TEC-ECN SIMULATION Q. Wijnands, TEC-SWM GS & OPERATIONS D. Patterson, OPS-OSA STRUCTURES H. v. de Graaf, TEC-MCS MISSION ANALYSIS OPTICAL METROLOGY RF METROLOGY G. Janin, OPS-GA A. Boutonnet, OPS-GA Z. Sodnik, TEC-MMO K. Wallace, TEC-MMO A. Garcia, TEC-ETN SYSTEMS THERMAL S. Santandrea, TEC-SP S. Gidlund, TEC-SYE G. P. Zoppo, TEC-SYE G. Chirulli, TEC-MCT Under the responsibility of: F. Teston, TEC-SP Study Manager P. Vuilleumier, TEC-SP Deputy Study Manager With the scientific assistance of: J-L. Boit, Laboratoire d'astrophysique de Marseille Technical consultancy provided by: J-M. Lautier, TEC-MMM Mechanisms S. Mangunsong, TEC-MCS Configuration The editing and compilation of this report has been provided by: S. Gidlund, TEC-SYE A. Pickering, TEC-SYE

6 page 4 of 231 This study is based on the ESA CDF Integration Design Model (IDM), which is copyright 2004 by ESA. All rights reserved. Further information and/or additional copies of the report can be requested from: F. Teston ESA/ESTEC/TEC-SP Postbus AG Noordwijk The Netherlands Tel: +31-(0) Fax: +31-(0) Frederic.Teston@esa.int For further information on the Concurrent Design Facility (CDF) please contact: M. Bandecchi ESA/ESTEC/TEC-SYE Postbus AG Noordwijk The Netherlands Tel: +31-(0) Fax: +31-(0) Massimo.Bandecchi@esa.int

7 page 5 of 231 TABLE OF CONTENTS 1 INTRODUCTION Background Scope Document Structure EXECUTIVE SUMMARY Study Flow Study Objectives Design Strategy Mission and Satellite Design for the Two Satellites MISSION OBJECTIVES AND REQUIREMENTS Background Mission Objectives Mission Requirements Mission and Programmatic Constraints MISSION ANALYSIS Orbit Trade-Off Baseline Orbit Orbit type Selection of the Orbit parameters Orbit maintenance Launchers Performance Launch Window Ground Coverage Formation Initiation Formation Keeping Assumptions Coronagraph Science Mode: Perigee pass Coronagraph Science Mode: Sun Pointing FF Demonstration manoeuvres End of Life De-orbiting SYSTEMS Mission Requirements and Constraints Orbit trade-off Launcher trade-off Launcher Propulsion Module Number of Satellites trade-off Orbit Utilisation and Segmentation...35

8 page 6 of Solar corona observations FF demonstration manoeuvres Orbit segmentation Propulsion System trade-off System level trade off results Mass budget Two satellites Solution (Baseline) RF METROLOGY Requirements and Design Drivers Assumptions and Trade-Offs for the RF Metrology Subsystem Assumptions RF measurements error budget Link budget improvements for (Range vs. TM/TC data rate) RF metrology data budget analysis RF metrology antenna trade-off GPS requirements and manufacturers selection Two Satellites Solution (Baseline) RF metrology subsystem unit description RF metrology navigation performances IFs between RF metrology and, OBDH and TT&C subsystem RF metrology S-band antennas RF metrology open points and critical issues GPS unit description Recommended requirements for the RF metrology subsystem demonstration List of Equipment Configuration Options RF Metrology for the three satellites option GPS optional manufacturer OPTICAL METROLOGY Requirements and Design Drivers Assumptions and Trade-Offs Coarse lateral sensor (CLS) Fine lateral sensor (FLS) Absolute longitudinal sensor Fringe tracking sensor (FTS) Two Satellites Solution (Baseline) Coronagraph satellite Occulter satellite Options Three Satellites Option SUN OBSERVATION PAYLOAD (CORONAGRAPH) Introduction...69

9 page 7 of Coronagraph Science Objectives System Requirements and Design Drivers Coronagraph Design Overview GUIDANCE, NAVIGATION AND CONTROL GNC Main Modes Orbital Dynamics Orbital data Coordinates frames Relative dynamics Manoeuvres FF demonstration manoeuvres Coronagraph Science manoeuvres GNC Requirements V Computations Assumptions Station Keeping at 25m x-axis Station Keeping at 25m y-axis Station Keeping at 25m z-axis Station Keeping at 250m x-axis Station Keeping at 250m y-axis Station Keeping at 250m z-axis Station Keeping at 8km x-axis Station Keeping at 25m x-axis Standby Resize m along x-axis Resize m along y-axis Resize m along z-axis Resize m along x-axis Resize 1-8km along x-axis Rotation of 120deg in xz-plane at 25m Rotation of 120deg in xy-plane at 25m Rotation of 120deg in yz-plane at 25m Rotation of 120deg in xz-plane at 250m Rotation of 120deg in xy-plane at 250m Rotation of 120deg in yz-plane at 250m Rotation/Resize of 120deg in xz-plane at m Rotation/Resize of 120deg in xy-plane at m Rotation/Resize of 120deg in yz-plane at m Station Keeping Emma at 1km x-axis Coronagraph Science observation Pulse Perigee manoeuvre Sequencing Summary of V computations Number of Satellites Three satellites option Two satellites solution (Baseline)...86

10 page 8 of Relation to other European missions Demonstration coverage of customer missions Avionics Equipment for FF Propulsion and pulse width modulation Star tracker Gyros Coarse Sun sensor RF metrology Optical metrology CONFIGURATION Requirements and Design Drivers Assumptions and Trade-Offs Two Satellites Solution (Baseline) Coronagraph satellite Occulter satellite PROPULSION Requirements and Design Drivers Assumptions and Trade-Offs Electric propulsion trade off Cold Gas thrusters trade-off TRL Two Satellites Solution (Baseline) Coronagraph satellite Occulter satellite Options FEEP subsystem Three Satellites Option POWER Requirements and Design Drivers Mission overall requirements Coronagraph Science Orbit FF demonstration Orbit Power requirement on the bus Design methodology Assumptions and Trade-Offs Power architecture study Two Satellites Solution (Baseline) Budgets and power system configurations Coronagraph satellite Occulter satellite Options STRUCTURES Requirements and Design Drivers...137

11 page 9 of Assumptions and Trade-Offs Two satellites solution (Baseline) Coronagraph satellite Occulter Satellite THERMAL Requirements and Design Drivers Assumptions Two Satellites Solution (Baseline) Coronagraph satellite Occulter satellite TELECOMMUNICATIONS Requirements and Design Drivers Assumptions and Trade-Offs Data transmission assumptions Antenna trade-off Frequency band selection Two Satellites Solution (Baseline) TT&C architecture On-board TT&C system Modulations and coding On-board antenna selection and location On-board S-band transceiver Link budgets RF metrology Options List of Equipment Three Satellites Option DATA HANDLING SYSTEM Requirements and Design Drivers Assumptions and Trade-Offs Two Satellites Solution (Baseline) Three Satellites Option RISK General Approach Definition of the Safety and of the Mission Success Criteria Safety Mission success criteria Requirements Proposed for Mission Success Criteria, M1 to M Identification of Undesirable Events and Relevant Severity Categories The approach for mitigating the Undesirable Events occurrence Preliminary Approach for the Implementation of Risk Related Design Criteria Two satellites solution (baseline)...167

12 page 10 of Three satellites option SIMULATION AND VISUALISATION Introduction Requirements on the FES Description of the FES Simulation Results Ground station visibility D Visualisation GROUND SEGMENT AND OPERATIONS Requirements and Design Drivers Assumptions and Trade-Offs Satellite design and development Operations implementation LEOP recommendations Option of beacon monitoring Two satellites solution (Baseline) Routine operations support Mission phases Ground stations and communications network The mission control centre Computer facilities The flight control software system Mission Operations Concept Overview Mission planning, satellite monitoring and control Orbit and attitude control PROGRAMMATICS/AIV Requirements and Design Drivers Assumptions and Trade-Offs Two Satellites Solution (Baseline) Master schedule Model philosophy Model and test matrix AIV plan Summary and critical issues Three Satellites Option Master schedule (3 satellites option) Model philosophy Model and test matrix AIV plan Summary and critical issues COST...195

13 page 11 of Requirements Assumptions Organization Cost Breakdown Structure (CBS) Project environment Two Satellites Solution (Baseline) Baseline cost estimate Cost-risk analysis Three Satellites Option CONCLUSIONS Compliance with Study Objectives Compliance with Mission Objectives Compliance with Mission and Programmatics Constraints Overall Conclusions and Recommendations REFERENCES ACRONYMS APPENDIX A - GENERIC RF METROLOGY DESIGN AND DEVELOPMENT STATUS223 A-1 RF Metrology Subsystem Development Background Information A-2 RF Metrology Development Status A-3 Reference Frames A-4 Generic RF Metrology Communications Description and Link Budget A-5 RF Metrology Data Budget Analysis A-5-1 Interface (IF) between RF metrology and GNC A-5-2 IF between generated RF navigation data and TT&C...231

14 page 12 of 231 This Page Intentionally Blank This page intentionally left blank

15 Page 13 of INTRODUCTION 1.1 Background PROBA is a programme managed by the Technical and Quality Management directorate (D/TEC) initiated in 1998 for the in-orbit demonstration of platform and payload technologies for potential application to future missions. PROBA1 was launched in October 2001 with the aim to demonstrate on-board autonomy (see PROBA2 was initiated in 2004 with the aim to validate hardware and software advanced space technologies. is dedicated to the demonstration of Formation Flying (FF) technologies, currently under development in Europe, in preparation for future European Scientific and application missions. Following an industrial study performed on the subject in 2004, an ESA internal study was requested by the Special Programme office of D/TEC using the resources of the ESTEC Concurrent Design Facility (CDF). The purpose of the study, funded by the ESA General Studies Programme (GSP), was to assess the feasibility of an affordable demonstration mission, accommodating an instrument dependent on FF mission characteristics. The selected instrument is a coronagraph, used for Sun-corona observation. In the frame of the ESA/CNES cooperation, the French Space Agency started a similar study, named ASPICS2, in the same period at their PASO Centre d Ingénierie Concourrante (CIC) in Toulouse. ESA and CNES specialists participated in the other team s sessions, and the final study results were shared and compared. 1.2 Scope The objectives of the study were to perform mission conceptual design and trade offs; prepare a preliminary design including budgets, configuration, subsystem designs with the required performances, define space and ground operations, define programmatics, perform risk assessment and cost analysis. The study also investigated what was achievable in terms of technological demonstration within the frame of the mission. Finally, this study serves as preparation and input for the upcoming industrial phase A. 1.3 Document Structure The Executive Summary contained in the next chapter provides an overview of the study flow and objectives, design strategy and mission and satellite design. Subsequent chapters provide details on each domain addressed in the study. Due to the different distribution requirements only cost assumptions excluding figures are given in this report. The costing information is published in a separate document, CDF-42(B).

16 Page 14 of 231 This Page Intentionally Blank

17 Page 15 of EXECUTIVE SUMMARY 2.1 Study Flow The study consisted of eight sessions, beginning with a kick-off on the 9 th September 2005 and finishing with an Internal Final Presentation (IFP) on the 11 th October Following the IFP, additional off-line work was conducted on specific topics, involving a restricted complement of the study team. In particular this phase of the study was dedicated to the optimisation of the delta-v usage throughout the mission, in order to maximise the mission duration. The results were then consolidated with all the members of the interdisciplinary team in a conclusive session on the 8 th. 2.2 Study Objectives The study objectives were to: Review the Formation Flying (FF) technologies requiring in orbit demonstration and the related requirements Assess the technical feasibility of a FF demonstration mission fulfilling the payload requirements within the given constraints Identify the best compromise between the number of, and the level of, demonstrated FF technologies and mission complexity, cost, and the interest for future ESA missions Identify the derived mission and system requirements Present a preliminary system baseline design, including trade-offs, analysis and the identification of critical aspects and areas Present a preliminary FF mission implementation plan and cost envelope identification For the de-scoped FF technologies, provide an indication of the rationale and their impact on system design. 2.3 Design Strategy The first iteration aimed at fulfilling the full set of technology demonstration objectives, but due to various mission constraints some of the objectives had to be de-scoped in the following design iterations. Trade offs have been performed on the number of satellites required, operational orbit, launchers and propulsion systems. The resulting baseline design includes two satellites with ample mass and on-board resources margins. The study also provides recommendations and design options to be further investigated for the incorporation of some of the discarded mission objectives. Specifically the three satellite FF mission has been studied during the sessions, in order to study and understand the related requirements. 2.4 Mission and Satellite Design for the Two Satellites Table 2-1 provides a schematic summary of the mission and satellite design for the two satellite solution. For further details, see relevant chapters in this report.

18 Page 16 of 231 Mission objectives Launch Orbit Operations Occulter satellite In-flight demonstration and validation of FF technologies, currently under development in Europe, in preparation for future European scientific and applicative missions: o Metrology systems (RF Metrology, Optical Metrology); o Actuators (High precision propulsion systems); o FF System architectures (Centralised, Decentralised); o FF GNC design; o Typical FF manoeuvres; o Validation of engineering tools, simulators, test beds; o End to end command & control and operations. Consolidation of the engineering approach required by a FF mission including new verification and validation techniques, Accommodation of a Payload, Sun observation coronagraph Mission duration 1 to 2 years Mid Launcher: VEGA/VERTA or Eurockot/DNEPR (both options include Lisa Pathfinder Propulsion Module) Orbit type 24h HEO Perigee altitude 800km Apogee altitude 70772km Eccentricity Period 86164s = 23h 56mn 4s Inclination 60 VEGA/VERTA (63 with Eurockot/DNEPR) RAAN Depending on launch window Eclipses Maximum 1.4h Ground station Perth Visibility FF demonstration manoeuvres Design lifetime Attitude control Mass Satellite main body dimensions Propulsion Power 20h/day Minimum of 12 days including: Station Keeping at various Inter-Satellite Distances Forced motion along a line In plane fly around Simultaneous rotation and resize 19 months Three-axis stabilised, 6 degrees of freedom (DOF) Dry mass 210.9kg Wet mass 236.4kg 0.756m x 0.850m x 0.900m Disc with 1.2m diameter mounted in Sun direction Cold Gas, 12x10mN & 12x40mN thrusters 25kg propellant (N 2 ) AsGa TJ solar array, 1.05m 2 disc) 30V regulated bus Li-Ion Battery, 701.2Wh (Body-mounted on

19 Page 17 of 231 Coronagraph satellite Communications Data handling Thermal RF metrology Optical metrology Coronagraph Payload Design lifetime Attitude control Mass Satellite main body dimensions Propulsion Power Communications Data handling Thermal RF metrology Optical metrology Coronagraph Payload Redundant transceiver module, 4 S-Band LGAs 1 processor module, 1 mass memory module Passive control 1 RF FF unit, 6 S-band LGAs Coarse lateral unit; Fine lateral sensor (FLS) receiver; Absolute longitudinal sensor (DWI); Fringe tracking metrology (OPD) Disc (external occulter) 19 months Three-axis stabilised, 6 DOF Dry mass 210.7kg Wet mass 218.6kg 0.756m x 0.850m x 0.900m Cold Gas, 12x10mN & 12x40mN thrusters 8kg propellant (N 2 ) AsGa TJ solar array, 1.5 m 2 (2 x panel 0.75 x 1.0m) 30V regulated bus Li-Ion Battery, 797.6Wh Redundant transceiver module, 4 S-Band LGAs 1 processor module, 2 mass memory modules Passive control 1 RF FF unit, 6 S-band LGAs 1 Corner cube reflector, 1 FLS transmitter Telescope and electrical box Table 2-1: Mission and satellite design characteristics

20 Page 18 of 231 This Page Intentionally Blank

21 Page 19 of MISSION OBJECTIVES AND REQUIREMENTS 3.1 Background FF missions are envisaged in different space domains to enable new missions that cannot be performed by conventional systems. For example, telescopes with focal lengths of several hundred meters. Examples of ESA future missions using FF are DARWIN and XEUS. The success of the envisaged missions depends on understanding and correctly implementing all the aspects of FF. For this purpose, in-flight demonstrations are proposed for FF, and required technologies. Technological developments enabling FF have already been initiated and some precursor missions have been investigated, for example SMART-2, SMART-2+, SMART-3. In addition, an interest in a FF demonstration mission has been expressed in the frame of the PROBA-2 Call-For-Ideas by several ESA member states. 3.2 Mission Objectives The primary mission objective, for the study, is to demonstrate FF through: Flight qualification of FF hardware, which currently are foreseen to be: o RF metrology o Optical metrology o High accuracy propulsion system o FF GNC system Demonstration of FF manoeuvres for future missions Exploitation of the FF for Scientific return. 3.3 Mission Requirements The primary objective of the mission, as described above, result in the following specifications for the two main topics of this mission, FF and coronagraph science: In-flight demonstration and validation of FF technologies currently under development in Europe, prior to their use by future ESA operational missions, including to the maximum feasible extent: o Metrology systems (GPS, RF Metrology, Optical Metrology and so on) o Actuators (high precision propulsion systems) o Inter Satellite Communication Links (ISL) o FF system architectures (Centralised, Decentralised, Distributed) o FF GNC design o Typical FF manoeuvres o Validation of engineering tools, simulators, test beds o End to end command and control and operations o Consolidation of the engineering approach required by a FF mission including new verification and validation techniques Accommodation of a Sun Observation Payload (Coronagraph)

22 Page 20 of Mission and Programmatic Constraints The following items present the mission and programmatic constraints, as provided by the study manager. The constraints provide the desired envelope within which the study solutions and results are to be presented: satellite development and launch shall take place in the near future, not later than 2010, in order to provide a timely feedback for operational ESA missions planned for the next decade. The Operational Phase shall be from 2010 to 2012 The mission shall be designed with a low cost approach and is assumed to cover: o System Analysis o Additional development of Formation Flying Technologies to achieve flight readiness o Ground Verification o Procurement of Platforms o Assembly, Integration and Test o Launch and Operations/Data Analysis GSTP shall provide the programmatic framework for the implementation of is planned for initiation in 2006 The implementation shall follow the classical Satellite Project Approach o Phase A: Feasibility including enabling technologies up to Technology Readiness Level (TRL) 4 o Phase B: System Definition including technologies pre-development up to TRL 6 o Phase C/D: System Procurement o Phase E: System Operation

23 Page 21 of MISSION ANALYSIS 4.1 Orbit Trade-Off Various types of orbits can be contemplated for the PROBA 3 mission: Low Earth Orbit (LEO), for instance of Sun-synchronous type, popular for remote sensing missions Geostationary Orbit (GEO), ideal for full ground coverage from a single station High Eccentricity Orbit (HEO), ideally its period should be synchronous to the Earth rotation with a period of 12-, 24-, 48-h or higher, popular in past ESA astronomy missions Libration Point Orbits (LPO) around Sun-Earth Lagrange point L1 or L2, the place of choice for missions requiring a quiet environment. Table 4-1 lists the pros and contra of these orbits. Orbit Pro Contra Low launch energy High disturbances (e.g. drag, gravity LEO Short communication distance potential, South Atlantic anomaly) Frequent eclipses Short ground coverage periods ( 8 mn/pass) SSO Like LEO + Nearly-constant solar incidence on orbital plane No eclipse when h > 1382 km Complete coverage with one GEO station with non-steerable dish Good and regular coverage HEO (synchronicity) Eclipse reasonably short (launch window) LPO Quiet environment No eclipse (Halo orbit) Constant distance from Earth and Sun Like LEO High injection requirement (1.5 km/s) Orbit maintenance (50 m/s per year) Orbit unstable (launch window) Large communication distance at apogee Orbit synchronicity maintenance (1 m/s per year) High launch energy requirement Orbit maintenance (0.2 m/s per year) Very large communication distance (up to 1.8 million km) Long transfer time ( days) Table 4-1. Pro and contra of popular terrestrial orbits The most suitable orbit for is an LPO. However, based on considerations on launch energy and communication distance, an HEO 24-h type was selected as more appropriate for a low-cost mission. Also, for a demonstration mission, FF in HEO is more challenging than in LPO.

24 Page 22 of Baseline Orbit Orbit type The orbit trade-off results in the selection of an HEO. To maximise ground coverage, a synchronous orbit is recommended. Synchronous HEOs have been quite popular among ESA projects and the main types of synchronous HEOs have been populated by satellites (12-h: GEOS, 24-h: ISO, 48-h: XMM, 72-h: INTEGRAL). Selected for PROBA 3 is a 24-h orbit, allowing very good daily ground coverage with only one station Selection of the Orbit parameters Orbit period: one sidereal day: s = 23 h 56 mn 4 s. Perigee height: the launcher will deliver the satellite with a low perigee between 180 and 300 km. In order not to be too sensitive to orbit perturbations, the perigee height needs to be raised. Two perturbations are playing a major role: Atmospheric drag during perigee passage: reduces the orbit energy, therefore the apogee height. Luni-solar gravity pull at apogee: changes the eccentricity, therefore the perigee height. If the perigee height decreases too much, the satellite will burn up in the atmosphere. Both effects need to be counteracted in order to keep the orbit synchronous. Taking these effects into account, a recommended minimum perigee altitude is 800 km (for INTEGRAL, 1000 km was selected). Apogee height: the 24-h orbit apogee height corresponding to an 800 km perigee height is km. Inclination: to have maximum ground coverage, a high inclination combined with a high latitude ground station is favoured. Payloads launched by Russian rockets are injected into orbits with an inclination in the area of 60. This is a good value for a synchronous HEO inclination. For a launch from a low latitude site, such as Kourou, targeting such a high inclination will penalise the launcher s performance. Argument of perigee: optimum coverage is obtained when the declination of the apogee is as high as possible. This is achieved with an argument of perigee of 90 for a ground station in the southern hemisphere, or 270 for a station in the northern hemisphere. Right Ascension of Ascending Node (RAAN): as mentioned above, due to luni-solar perturbations HEOs are unstable and the launch hour, therefore the RAAN, has to be selected in order to have the eccentricity initially decreasing (perigee raising). However, a mission constraint (see Section 4.7.2) imposes a RAAN in the vicinity of 0. The effect of this constraint will be shown in Section Orbit maintenance Regular orbit manoeuvres are needed for maintaining synchronicity. These manoeuvres are most efficiently performed at perigee and amount to about 1 m/s per year.

25 Page 23 of Launchers Performance The baseline launcher for PROBA 3 is VEGA/VERTA. Other contemplated launchers are Rockot and Dnepr. None of these launchers are able to put a substantial payload on the baseline operational orbit. It is therefore proposed to: Use the launcher only for reaching a LEO Equip the payload with its own propulsion unit, acting as an upper stage for raising the apogee height to its nominal value of km. Restart the propulsion unit at apogee to raise the perigee height to 800 km. This procedure is similar to the launch procedure of Lisa PathFinder (LPF), the project formally known as Smart-2, for reaching an orbit around libration point L2, where a Rockot launcher (baseline launcher for LPF up to the time when this report was written) injects the spacecraft into a km LEO. After separation, the spacecraft performs the injection into transfer orbit using a propulsion module (RD[1]). The propulsion system selected by LPF is the European bi-propellant 400N engine with a specific impulse of 321s and Eurostar tanks. As this engine has a rather low thrust level, the injection manoeuvre is performed in several perigee burns so that gravity loss is kept to a minimum. In the case of LPF, not less than 15 consecutive burns are scheduled. Table 4-2 shows, for launchers Rockot (RD[2]), Dnepr (RD[3]) and VEGA (RD[4]), the performance into Parking Orbit (PO), the dry mass of the satellite plus propulsion module, the type of Eurostar tank used, the total V needed for the propulsion unit to reach final orbit, the corresponding propellant load and the satellite useful mass. The baseline option (VEGA) has been highlighted. Rockot Dnepr VEGA (1) VEGA (2) PO 200x Inclination Perf. In PO Dry mass Tank type 2000 short 3000 large 2000 short 2000 long Total V Prop. mass Useful mass Table 4-2: Mass performance of launchers Rockot, Dnepr and VEGA for orbit acquisition Remarks concerning Table 4-2: Adapter mass is included in satellite dry mass Residual propellant and consumable are taken into account Gravity loss is assumed to be negligible (a multi-burn injection is assumed)

26 Page 24 of 231 The useful mass listed in the last row is the dry spacecraft mass, without propulsion unit (motor, tanks and structure) VEGA launch: o An elliptic PO similar to Rockot s PO for LPF would allow for a more efficient launch procedure. However, VEGA s performance in elliptic orbits is not known at this time. o The launch site (Kourou) is at a low latitude (5.4 ), which penalises performance for high inclination target orbit. For information, in the last column (VEGA (2)) mass figures are indicated for a low inclination case. 4.4 Launch Window For a nominal lifetime of 19 months, a launch window has been calculated along the year It is represented in Figure 4-1 in the form of level-lines in a day/hour diagram. The following parameters are represented: Orbit stability: in the region shaded by (in red), the perigee height decreases below 600 km during the lifetime of 570 days Eclipse duration: black level-lines show eclipse duration in hours RAAN: RAANs outside a band ( 20, 20 ) are shaded by + (in blue, level lines labelled 340, 360 and 380). Figure 4-1: PROBA 3 launch window diagram.

27 Page 25 of 231 Figure 4-1 shows that: The window is open in the blank region, limited on the bottom by the (blue) 340 RAAN constraint and on the top by the (red) minimum lifetime of 570 days Maximal eclipse duration in the open area is between 1.0 and 1.4h. It is possible to aim for a 0 RAAN the whole year. 4.5 Ground Coverage Baseline selected ground station for is Perth. With the selected synchronous 24h orbit, ground coverage from Perth is constantly 20.8 h/day for a 5 minimum elevation. Note: should a low inclination 24-h orbit be chosen, coverage would only be 15.6h/day for 10 inclination. 4.6 Formation Initiation The propulsion module, as well as the 2-satellite stack and the adapter, are released from the launcher just below the operational orbit in terms of apogee altitude. The first satellite (the Coronagraph) is separated with the help of a tangential manoeuvre along the velocity vector (approximately 30 cm/s at the perigee) putting it into the operational orbit. When the satellite crosses the propulsion module orbit one revolution later, the distance between them is about 850 km. The Occulter satellite is released one revolution after the Coronagraph, with the same manoeuvre. Consequently, the formation will drift away from the propulsion module, while the distance between the two satellites will remain bounded between 80 km and 850 km. The end of the initialisation of the formation is achieved by using a two tangential manoeuvres strategy with the Occulter satellite. The Delta-V consumption will be related to the duration (called ndays) of this last phase: its order of magnitude will be 60/ndays cm/s. As an example, the Delta-V will be 6 cm/s if the duration is assumed equal to 10 days. In order to insure no collision risk between the propulsion module and the formation of satellites, a last manoeuvre could be performed by the re-ignitable propulsion module to put it into a different orbit. 4.7 Formation Keeping Assumptions The following is assumed: Mission duration: 19 months Satellite mass of 200kg The ISD in the coronagraph science mode is 120m Only 2-body dynamics are considered, as depicted in Figure 4-2. This figure gives the ratio between the central potential acceleration and other perturbation accelerations, namely J2, Sun and Moon. This ratio is given as a function of the time after perigee pass. It is clear that all perturbations can be seen as noise when assessing the required thrust and V budget for the manoeuvres.

28 Page 26 of 231 One orbit is divided into two parts: the limit between them is a given true anomaly v. Between v and 2 π v, the formation is either in coronagraph science mode (Sun pointing), or in FF demonstration manoeuvres mode. Between 2 π v and v, the formation geometry is rearranged to save propellant mass. Ratio between central potential gradient and other gradients Time (hours) J2 Moon Sun Figure 4-2 Ratio between central potential acceleration and other perturbations acceleration Coronagraph Science Mode: Perigee pass As stated above, the formation is rearranged around the perigee to save propellant mass. At the true anomaly equal to 2 π v, a first manoeuvre is applied. After this, the satellites fly in free fall. Close to the true anomaly v, a second manoeuvre is applied to re-enter the coronagraph science mode. This principle is schematised in Figure 4-3. Out-ofplane angle DV1 In-plane angle DV2 Figure 4-3: Basic principle of formation keeping

29 Page 27 of 231 The duration of each manoeuvre is assumed to be one hour. The maximum thrust needed is given in Figure 4-4 as a function of the duration of the coronagraph science mode (that is, as a function of the true anomaly v ). For a 12-hour Sun observation, the corresponding thrust is around 0.4mN. Figure 4-4: Maximum thrust needed during perigee pass The annual V budget for this phase is given in Figure 4-5. It shall be noted that the thrust out of the thruster direction is not taken into account, but the V margins are ample enough to cover this issue. For a 12-hour Sun observation per orbit, the V budget over the mission duration is 7.5m/s. The last point to be analysed is the collision risk during the free fall. Let us define two important angles (see Figure 4-3): The observation in-plane angle: this angle is measured within the orbital plane of the satellite. It is counted from the direction of the perigee. The observation out-of-plane angle: is the angle between the direction of observation and the orbital plane of the satellite. It has been shown during the study that when the out-of-plane angle increases, the minimum distance decreases. It finally reaches 0m when the out-of-plane angle value is 90. As a consequence, an easy way to limit the collision risk is to keep the out-of-plane angle small. By choosing a RAAN equal to 0 the out-of-plane angle can be kept between and Any other value would lead to a higher out-of-plane angle.

30 Page 28 of 231 Figure 4-5: V budget for perigee path The minimum ISD is given in Figure 4-6. For a 12-hour observation, it corresponds to a minimum distance equal to 45m. Figure 4-6: Minimum ISD during perigee pass

31 Page 29 of Coronagraph Science Mode: Sun Pointing Figure 4-7 gives the maximum thrust needed during coronagraph science mode. For a 12-hour observation, 50µN is needed. This value corresponds to the worst case Sun direction over the mission. Figure 4-7: Maximum thrust needed during coronagraph science mode Figure 4-8 gives the annual V budget. For a 12-hour observation per orbit, the mission V budget is 6.3 m/s. Figure 4-8: V budget for the coronagraph science mode

32 Page 30 of FF Demonstration manoeuvres The V budget and thrust levels required for the FF demonstration manoeuvres can be found in the GNC Chapter. 4.8 End of Life De-orbiting The selected HEO for PROBA 3 is subjected to strong luni-solar perturbations and the perigee height is expected to decrease and reach the dense atmosphere within 20 years. The Space Debris requirement of leaving no object in near Earth orbit for more than 25 years is therefore met without the need of a de-orbiting manoeuvre. No extra V for spacecraft de-orbiting at end of operational life is required.

33 Page 31 of SYSTEMS 5.1 Mission Requirements and Constraints The following high level mission requirements were given at the beginning of the Study: To validate relevant FF technologies in flight, within the constraints of a small mission and limited budget To embark a user payload, to maximize the investment return of the mission To minimize the launch cost and follow a low cost development approach (involving an extensive use of COTS equipment for the platform, wherever possible) To target a mission launch date at the end of the decade, to provide a timely feedback for the development of ESA operational missions based on the use of FF. FF is an operational technique currently envisaged to be used in many different contexts, spanning, with its most diverse forms, from Earth Observation missions to Science missions. Several of the needed technologies are currently under development in Europe, but the full implementation of a FF mission is still to be done. Scientific missions, because of their highly demanding, relative position and attitude knowledge and control accuracy requirements, represent the natural principal target of a demonstration mission like since they encompass the mission requirements from other domains. This Feasibility Study therefore focuses more on mission and spacecraft design for demonstrating the low perturbation FF foreseen in future missions located in unbound orbits than on design for FF in LEO. The following table shows the conceptual differences between the two cases, and the studied mission design. EO Science Orbit LEO to Very LEO LPO 24h HEO Environment Very high perturbation levels, Earth Potential and Atmospheric Drag the largest contributors Type of FF FF manoeuvres Same ground-track, satellites distributed on the same orbital plane, Cartwheel orbits Drag compensation, FF Station-keeping, Resizing Very low perturbation levels Non-Keplerian forced formation FF Station-keeping, Resizing, Slewing, Rotations and combinations of the above manoeuvres Low perturbation levels at apogee, high levels at perigee One satellite in a Keplerian orbit, the other performing non-keplerian FF (FF only during apogee), also same ground-track FF Station-keeping, resizing, rotations, simultaneous rotation and resizing Table 5-1: Characteristics of FF for both EO and Science missions, compared with the FF approach

34 Page 32 of 231 The outcome of the orbit and launcher trade-offs determines the technology demonstration equipment embarked and the number of repetitions of the technology experiments sets, as well as the FF techniques and maneuvers that can be included in the baseline design. The final baseline design provides a margin for the incorporation of additional technology demonstration experiments. Examples of utilization of the remaining available onboard resources are given in the relevant Chapters. Beside the baseline design, ways to incorporate the de-scoped technologies and FF techniques were investigated. As an example, the incorporation of a limited set of Electrical Propulsion modules to be embarked, for in orbit validation, in addition to the platform propulsion system, has been investigated. The description of system level requirements specifically imposed by the scientific payload (coronagraph), the technological experiments and the system level impact of the FF techniques to be demonstrated is provided in the relevant Chapters. 5.2 Orbit trade-off The orbit trade-off led to the following conclusions: Lagrange Point Orbits (LPOs) would provide the best operational scenario for the demonstration of technologies related to the future ESA scientific missions. However, the spacecraft mass available in operational orbit is limited when small launchers are considered. The selection of this kind of orbit would cause the de-scoping of additional FF technology experiments and/or of the scientific payload. Low Earth Orbits (LEOs) provide the highest spacecraft mass into operational orbit. However, for formation flight, a typical microsatellite would not manage the level of perturbations in this orbit, therefore not allowing non-keplerian FF maintenance. Geostationary Earth Orbits (GEOs) are too expensive to reach with the given launchers and available resources and have been immediately traded off Highly Elliptical Orbits (HEOs)/Geostationary Transfer Orbits (GTOs) can allow FF technologies demonstration and coronagraph observations in a relatively quiet environment during the majority of the orbital period (around the apogee leg) while at the same time allowing acceptable overall spacecraft mass into operational orbit: o 12h GTOs could in principle benefit from the availability of piggy-back opportunities with commercial launches (imposing however constraints on satellite mass and envelope as well as the launch date/mission timeline) o 24h HEOs provide overall better conditions compared to the 12h GTOs (Sun illumination, radiation environment, low perturbation environment periods duration and quality around the apogee, etc.), but require a dedicated launch. The considered orbit options and their characteristics are summarised in Table 5-2.

35 Page 33 of 231 Table 5-2: Orbit options and their characteristics Based on the trade-off conclusions, the target nominal operational orbit selected for is a 24h HEO with the following characteristics: Perigee altitude: 800km (chosen in order to minimise the atmospheric drag perturbation component) Inclination: 60 with VEGA/VERTA launch (63 with Eurockot launch) Eccentricity: Visibility (typical): 20h/day from ESA Perth GS Eclipses (typical): from 0 to 1.3h (depending on the RAAN). 5.3 Launcher trade-off Launcher The following conclusions were the results of the launcher trade-off, for the targeted operational orbit: A dedicated launch using a small launcher with an additional propulsion module allows a cost-effective launch into the target orbit, if the additional propulsion module is already existing or under development Piggy back solutions, for example Ariane5/ASAP, were discarded, for the following main reasons: o Few, if any, opportunities for launch to the selected 24h HEO are available o In the case of a piggy back launch to a commercial GTO, a dedicated high thrust propulsion system is required onboard each satellite to raise the initial GTO to the desired 24h HEO. Due to launcher constraints on the envelope, stacking the satellites on top of a single propulsion module is not feasible. Remark: If on the other hand the target orbit is changed to a 12h GTO, piggy-back solutions can become available with commercial launches. However, the stringent mass and envelope constraints related with a piggy-back solution will introduce additional limiting factors on the satellite design. In addition, perigee raise manoeuvres will be required in order to achieve the operational 12h GTO from the initial GTO (having a perigee altitude in the 200 km range). This would require a dedicated propulsion system or module on each satellite of the formation.

36 Page 34 of 231 The performances of a range of launchers were investigated for the chosen 24h HEO option and based on this analysis, the following candidate and backup options were selected: 1. VEGA/VERTA, 2. Eurockot/Dnepr Propulsion Module An additional propulsion module will be needed with all the selected launch options in order to reach the operational orbit. Two options have been considered: 1. Lisa Path Finder Propulsion Module (LPF PM). Liquid, re-ignitable, including reaction control system, existing design of an interface to the VEGA launcher, currently under development at EADS for the Lisa Path Finder mission 2. STAR37 family Solid Rocket Motors. Solid motors, requiring ad-hoc implementation of structural and electrical interfaces and further module and reaction control system design effort. The LPF PM has been selected as baseline. 5.4 Number of Satellites trade-off The initial design iteration aimed to entirely fulfil both the technology demonstration and scientific return requirements of the study, as follows: The incorporation of the coronagraph instrument (telescope and electrical box on one satellite and a 1.2m diameter occulter disc on a second satellite) The incorporation of the full RF and Optical Metrology packages into the satellite platforms The realisation of a FF mission with three satellites for extensive FF manoeuvres demonstration (completely satisfying the demonstration requirements for the Darwin mission). The mass of this initial design was not compatible with a low cost launcher and therefore this option had to be discarded. Two alternative design solutions, one with three satellites and one with two satellites have been investigated, both retaining the coronagraph onboard. The three satellite solution requires a dramatic downsizing of the Optical Metrology (OM) package, with the exclusion of metrology modules at the high-end of the OM chain, thereby lowering the accuracy achievable in the frame of the FF demonstration. The two satellite solution retains the full optical metrology package, but reduces the FF demonstration. The two satellites solution results in a total mass of 455kg (505kg including 50kg for the interface to the LPF PM), which is well within the VEGA/VERTA launch capabilities. Therefore this option was chosen as the baseline. For details on the mass budget of this option, see Section On the other hand, the overall mass of the three satellites option remained above the considered launchers capabilities even after the optical metrology package had been sensibly reduced. Because of the limited timeframe, this option was not studied in detail and optimised. In the subsystem chapters the changes in design needed for a three satellites option are described,

37 Page 35 of 231 whenever necessary, as well as the advantages this option would provide to the overall mission return. In this Report, the satellite embarking the coronagraph optical bench and electronic box is always referred as the Coronagraph or Coronagraph satellite, while the terms Occulter and Occulter satellite are used to refer to the spacecraft carrying the occulter disk. In the three satellite option, the third satellite, whose design has been assumed to be a replica of the Occulter, with minor differences in the embarked OM elements only, is often referred to as Free Flyer. 5.5 Orbit Utilisation and Segmentation Solar corona observations The vast majority of the mission operational orbits will be used for the operational FF required for the observation of the Sun-corona with the embarked scientific payload. On each of these orbits, observation of the Sun will be carried out over a 12 hours long orbital segment centred across the orbit s apogee, with an instrument s maximum on-time envisaged to be equal to six hours within this 12 hours period. During the complete orbital segment allocated to Solar Corona observations (12 hours), the Occulter satellite is to be kept at a constant distance of 120m from the Coronagraph satellite, in the Sun direction. The relative position and pointing control requirements for this operational phase are reported in Chapter 8. To achieve the absolute and relative attitude control requirements of the coronagraph, a set of four mini reaction wheels has been envisaged for both satellites, as a complement to the set of actuators (thrusters) to be used for the FF demonstration manoeuvres. This approach reduces the impact of the Sun observation operational phase on the mission propellant budget FF demonstration manoeuvres A full set of FF demonstration manoeuvres is performed in six orbits and the delta-v cost will be approximately 19m/s. A minimum of two full set repetitions has been envisaged for the mission, with additional partial, or complete set repetitions dependant upon the remaining onboard resources (propellant). In order to minimise the impact of the FF experiments on the propellant budget, multiple manoeuvres are envisaged to be performed sequentially on the same orbit, in a 20 hours long segment centred around the orbit apogee. Further optimisation of this issue shall be targeted in the next design phases. For further details on the manoeuvres, the relevant delta-v breakdown and the considered avionic equipment, refer to Chapter Orbit segmentation Initially the calculation of the delta-v related to the operational phase of the mission included full formation active control along the whole orbit. With this approach, the satellites of the formation were maintained at a constant inter-satellite distance (ISD), during the perigee passes for both the Sun observation and FF demonstration orbits. This was an extremely conservative approach, which would only allow for two months of total mission duration, with typical microsatellite onboard resources (propellant mass).

38 Page 36 of 231 To overcome the high delta-v cost described in the above paragraph, each orbit has been segmented into four parts, with active control limited to the apogee part of the orbit, as described below: Apogee passage: During this orbital segment, either FF demonstration manoeuvres and technological experiments or Solar corona observations will be performed. During this orbit leg, the satellite formation shall be actively controlled. During Sun observations, both the satellites will be Sun pointed (with the Coronagraph behind the Occulter, with regards to the Sun). The attitude of the satellites during FF demonstration manoeuvres will depend on the requirements set by the manoeuvres themselves, as far as compatible with the overall system requirements. Breaking the formation: On each orbit, at the end of the apogee passage phase, to prepare the formation for the next perigee pass, the Occulter is manoeuvred from its current position into a slightly different Keplerian orbit, optimally selected to minimise the formation building/breaking propellant consumption. During this phase, the formation is actively controlled, and the non-manoeuvring Coronagraph is envisaged to be maintained Sun pointing. The attitude of the manoeuvring Occulter will instead be dependent on the direction of the thrust vector generated by the onboard propulsion system. A maximum of one hour per orbit has been allocated for this manoeuvring phase. Perigee passage: In this orbit phase, the two satellites are kept on two slightly different and properly selected Keplerian orbits, with collision avoidance maintained active for safety reasons. At this design stage, the two satellites are to be kept Sun pointed (unless collision avoidance manoeuvres become necessary) during this orbital leg. Four hours per orbit have been allocated for perigee passage for the FF demonstration orbits and ten hours for the orbits dedicated to Sun observation. Building the formation: At each orbit, at the end of the perigee passage phase, active control of the satellites is reactivated, and orbital manoeuvres are performed to position the Occulter satellite to the non-keplerian orbit required by the FF operations (Sun observations or FF demonstrations). During this phase, the non-manoeuvring Coronagraph is envisaged to be maintained Sun pointing. The attitude of the manoeuvring Occulter will instead be dependent on the direction of the thrust vector generated by the onboard propulsion system. A maximum of one hour per orbit has been allocated for this manoeuvring phase. Free-Flight Keplerian orbits Manoeuvres 2x1h Controlled Non-Keplerian orbits Figure 5-1: Orbit usage during Coronagraph science

39 Page 37 of 231 Free-Flight Keplerian orbits Controlled Non-Keplerian orbits Figure 5-2: Orbit usage during FF demonstration orbits This approach enables a mission duration of at least one year, and 19 months was chosen as baseline. The sizing factor becomes the duration of the apogee part for the orbits dedicated to Solar corona observation. However, the difference between an 8h and a 12h apogee passage is only 3m/s for a one year mission. Further details on the V-costs and calculations can be found in Chapter 4 and in Chapter 9. In Table 5-3 the two different approaches discussed above are compared. In the table the first approach has been extended to a mission duration of half a year to ease the comparison with the later (baseline) approach. Characteristics PRO CON Full orbit active control (First approach) Conservative approach. The S/Cs stay sun pointed with a constant separation distance of 100m the full orbit. Coronagraph science duration depending only on eclipse. FF manoeuvres during apogee The Inter Satellite Distance (ISD) is controlled throughout the orbit - very low collision risk. Very high coronagraph science return Very expensive in terms of Delta-V Partial orbit active control (12h apogee) Every orbit is split into four parts; one apogee part with FF manoeuvres or coronagraph science, one part before and one part after the apogee with manoeuvres to build/break the formation and a perigee passage part where the S/Cs are in anti-collision mode. Very cheap in terms of Delta-V During the perigee passage the S/Cs are left 'free-flying', further analysis needed in the future to completely cover all possible cases. (Higher collision risk than the Conservative approach) Partial orbit active control (8h apogee) Same as for 12h. Same as for 12h. Same as for 12h. Mission duration 6 months 12 months 12 months Duration of coronagraph science per orbit up to 24h (depending on eclipse) Delta-V 102m/s 9.5m/s 6.5m/s 12h Table 5-3: Overview of the different approaches to calculate the delta-v for the operational phase (Sun observation case) of the mission. 8h

40 Page 38 of Propulsion System trade-off The different levels of perturbations encountered in the environment of the selected operational orbit and the different required manoeuvres imposed severe and contrasting requirements on the required thrust range and propulsion subsystem. On one hand, a sufficiently high thrust level (in the mn thrust range) was required, while, on the other hand, the accuracy required by the FF demonstration manoeuvres required six degrees of freedom and a very fine control. Therefore, twelve thrusters have to be accommodated onboard each satellite, capable of providing a very low minimum impulse bit, to have a sufficient thrust resolution. As a result of the performed analyses, the required thrust range and resolution have shown not to be compatible with the performances of a single propulsion system. Given the mass and volume envelope of the satellites, it was not possible to accommodate two different EP systems or a hybrid EP/Cold Gas solution. The adopted baseline is instead based on a single micro-cold Gas system with two different sets of nozzles, to provide two different thrust levels, resulting in the necessity to accommodate 24 thrusters. However, the second set of nozzles increases the overall propulsion subsystem mass by only 1kg. The design of the onboard propulsion system shall however be revisited in the following mission design phases, in view of the future operational missions requirements and design. 5.7 System level trade off results Figure 5-3 below summarises the trade-offs as described in the previous Sections, highlighting the retained solutions. LEO GTO HEO Orbit Lagrange Points Piggy back small launcher PM + Dedicated small launcher PM + Piggy back heavy launcher Dedicated medium launcher Launcher 2 S/Cs 2 S/Cs 3 S/Cs 2 S/Cs 3 S/Cs 2 S/Cs 3 S/Cs # of S/Cs Each S/C can have one of the following Propulsion CGS EPS EPS N 2 H 4 N 2 H 4 N 2 H 4 CGS CGS EPS Figure 5-3: Option tree with retained solution in green.

41 Page 39 of Mass budget The system mass budget is based on the inputs from the subsystems. On subsystem level a margin has been applied to the mass of units/items included, depending on the maturity status of the unit/item: 5% for off-the-shelf items 10% for items to be modified 20% for items to be developed. On system level a margin of 20% is added to the satellite dry mass, that is, the dry mass including subsystem margin. The propellant mass is added with a 2% margin, since the propellant tank cannot be completely emptied. No adapter between the LPF PM and the stack of satellites has been included in the system mass budget. It is estimated that the mass of this adapter will be around 50kg Two satellites Solution (Baseline) The total wet mass for the two satellites option is kg. Harness mass has been estimated to be 10% of the dry mass of the satellite, based on the harness mass of former PROBA satellites. The Occulter satellite system mass budget is shown in Table 5-4. Occulter Without Margin Margin Total % of Total Dry mass contributions % kg kg Structure kg Thermal Control 4.07 kg Communications kg Data Handling kg GNC 7.59 kg Propulsion kg Power kg Harness kg Instruments kg Total Dry(excl.adapter) kg System margin (excl.adapter) % kg Total Dry with margin (excl.adapter) kg Other contributions Wet mass contributions Propellant kg Adapter mass (including sep. mech.), kg 0.00 kg Total wet mass (excl.adapter) kg Launch mass (including adapter) kg Table 5-4: System mass budget for the Occulter satellite

42 Page 40 of 231 The Coronagraph satellite system mass budget is shown in Table 5-5. Coronagraph Without Margin Margin Total % of Total Dry mass contributions % kg kg Structure kg Thermal Control 3.92 kg Mechanisms 3.20 kg Communications kg Data Handling kg GNC 7.59 kg Propulsion kg Power kg Harness kg Instruments kg Total Dry(excl.adapter) kg System margin (excl.adapter) % kg Total Dry with margin (excl.adapter) kg Other contributions Wet mass contributions Propellant 7.80 kg Adapter mass (including sep. mech.), kg 0.00 kg Total wet mass (excl.adapter) kg Launch mass (including adapter) kg Table 5-5: System mass budget for the Coronagraph satellite

43 Page 41 of RF METROLOGY 6.1 Requirements and Design Drivers The RF FF metrology subsystem is the first element in the metrology system chain (together with the sun sensor and the star tracker). These sensors will ensure initial good relative attitude and position accuracy for the subsequent optical metrology systems (coarse lateral metrology, fine lateral metrology and fine longitudinal metrology). The RF FF metrology subsystem provides autonomous restitution of coarse relative position and attitude (option), throughout all the mission phases. The RF metrology subsystem computes the relative navigation between the satellites for the GNC subsystem. The main requirements for the RF FF metrology subsystem are: RF metrology is one of the developments to be tested in space as part of the mission objectives Computation of relative navigation: position, attitude (option), velocity, attitude rate (option) and local time, for all satellites in the formation, in all operational phases of the mission; with two basic functions, collision avoidance and coarse position. Provision of coarse positioning to acquire the optical metrology from the RF metrology, requiring 1 cm longitudinal accuracy and +/- 1 lateral accuracy. During the deployment, it is used for anti-collision and to set up the nominal formation. During the nominal formation, it is used for anti-collision and during formation reconfigurations. Provision of a local inter-satellite data link (TM/TC) of 9 kbps for the Master (Coronagraph) and 3 kbps for the Free-flyers. Provision of a synchronized clock reference (with the rest of the satellites) to each satellite (Local System Time). Autonomy and robustness. Full sphere visibility (4π steradian). Hybridization is required with other attitude sensors (for instance, star trackers) to provide full relative attitude and attitude rate restitution with respect to the body related frame (and, eventually, with respect to an inertial frame). In the case of three satellites no hybridization with other sensors is required. Use of GPS during the perigee passage (below 5000 Km) in order to synchronise the OBDH subsystem of the satellites and to provide absolute positioning for each satellite in the formation (to check the anti-collision functionality and the correctness of the RF relative navigation) Power consumption and mass should be minimised in order to maximise the mission duration Design should be kept as simple as possible in order to maximize the mission duration and to reduce cost. Data rate generation for on-ground experiments shall be optimised by giving realistic assumptions for the on-board equipment, TT&C and ground segment availability.

44 Page 42 of Assumptions and Trade-Offs for the RF Metrology Subsystem The generic RF metrology design and development status can be found in Appendix A. In this chapter the assumptions and adaptations for the mission application case are described Assumptions Formation configuration: The reference array configurations assumed for are: 2 satellites nominal configuration (baseline): it consists of two satellites (Coronagraph and Occulter) with the Centre of Mass placed in a straight line. 3 Satellites configuration (option): it consists of three satellites (Coronagraph, Occulter and a 3 rd satellite for FF demonstration) located in a plane, forming a triangle. Deployment configuration: random position and orientation of the satellites within a sphere of radius 8 km Assumptions during the deployment mode: The minimum distance between satellites (required for RF metrology acquisition) during the deployment phase is 5m and the maximum distance is 8 km (with data demodulation, and 15 km with only ranging function) The relative angular rate during deployment and nominal phases is < 5 deg/s Random position and orientation of the satellites within a sphere of radius r=8 km during deployment, with 0 < El <90 and 0 < Az < Assumptions during the nominal mode: The distance between the satellites during the nominal phase is between 25m and 250m. Any orientation of the satellites possible (any relative attitude) during nominal phase, with 0 < El <80 and 0 < Az <75 and 105 < Az <180. The relative angular rate during nominal phase is < 5 deg/s The maximum relative speed between satellites during the nominal phase is less than 0.5 m/s Operational assumptions: The FF RF subsystem navigation processing is decentralised during deployment, where each satellite computes the relative position to the rest of the satellites in the formation one by one (equivalent to a case of two satellites) The FF RF subsystem navigation processing is centralised during nominal phase, where one satellite computes the relative position of the satellites in the formation. In the centralised mode operation, it has to be demonstrated the tolerability to satellite failure (in the case of 3 satellites), with navigation accuracy performance degradation. At least 3 RF antennas are located in a plane perpendicular to the +X axis (3 antennas in each opposite side of the satellite), without RF obstructions.

45 Page 43 of RF measurements error budget The receiver pseudorange and carrier phase error were tested and validated with the FF RF subsystem BB. The error components are listed in Table 6-1 (RD[6]). However, the multipath error and biases are theoretical assumptions (maximum boundaries). These two elements are key parameters that need to be demonstrated. The antenna baseline for the angular measurements was assumed to be 1m. Measurement accuracy (1-σ) Pseudorange (P1/P2) 1.2 m Carrier phase (S1/S2) SD Carrier Phase (S1/S2) Mpath error (in P1 or P2) Mpath error (in S1 or S2) 2.5 deg 5 deg < 3m < 6 deg Chann frq antenna Biases (rms) 1mm 1 mm 1 mm Time tag accuracy 0.1 ms Table 6-1: Measured RF signal error budget with the FF RF metrology BB Link budget improvements for (Range vs. TM/TC data rate) This section describes the adaptation and improvements applicable to the case, starting from the generic RF metrology system described in Appendix A-4. There are three reasons why it is possible to increase the range or the TM/TC data rate in, use of only two sets of three antennas (instead of three sets), extra margin for the Free Flyers in the link budget, and less number of satellites (two (baseline) or three (option) satellites instead of four): uses only 2 sets of 3 antennas, then extra gain (2 sets): 2.2dB o Antenna set slot (2 sets): 7.5ms o Useful antenna slot: 5ms. 2.5ms are lost in acquisition and mis-syncronisation. Free Flyers extra gain: TM/TC data rate increment TM/TC increment for a fixed maximum range: 8Km TM/TC data rate due to extra gain: o Gain Master: 2.2dB (2 sets) 16Kbps TM/TC o Gain Free Flyers: 2.2dB (2 sets) (extra margin) 14Kbps

46 Page 44 of 231 The potential TM/TC increment is summarized in Table 6-2 TM/TC data 2 satellites (1 Master + 1 Free Flyer) Master: 14Kbps Free Flyers: 14Kbps 3 satellites (1 Master + 2 Free Flyers) Master: 16Kbps Free Flyers: 8Kbps Table 6-2: TM/TC data rate increment for a fixed maximum distance of 8 Km in Maximum range increment Maximum range increment for a fixed TM/TC data rate of: Master: n x 3 Kbps (n = 2, 3 in ) Free Flyer: 3 Kbps per satellite Table 6-3 provides the details and the conclusion of the analysis. Master Free Flyer 2 satellites (1 Master + 1 Free Flyer) Extra G: (fix rate) R: 32 km Extra G: (extra margin) R: 30 km 3 satellites (1 Master + 2 Free Flyers) Extra G: (fix rate) R: 18 km Extra G: (extra margin) R: 30 km Conclusion Rdata: 30 km limited by Free Flyer Rranging: 56 km Rdata: 18 km limited by Master Rranging: 31 km Conclusion Table 6-3: Maximum range increment for a fixed TM/TC data rate Since in both satellites have direct link with the Earth, and for the case of three satellites they use local wireless connection, it is recommended to fix the TM/TC data rate to: TM/TC: 9 Kbps for the Master (Coronagraph) (3Kbps x 3 additional satellites (Free Flyers)); and TM/TC: 3 Kbps for the Free Flyers, with a maximum ISD of 30km for the case of 2 satellites and 18 Km in case of three satellites RF metrology data budget analysis Interface (IF) between RF metrology and GNC The RF metrology interacts with the GNC subsystem in order to provide the coarse navigation and attitude estimations at 1Hz

47 Page 45 of 231 The communication between the RF metrology and the GNC subsystems is done via the OBDH at 1Hz. Both subsystems are located in the same satellite The data interchanged depends on the FF status, centralised (Coronagraph) in nominal mode and decentralised in deployment mode. Table A-6 in appendix A provides the RF metrology data description of the RF IF to the GNC subsystem, illustrated only for 1 satellite (see RD[8]). RF metrology data to the GNC subsystem: Centralised in the Coronagraph (no info about the Coronagraph coordinates needs to be sent, since it is the reference frame (RBF)): o 2 satellites: 832bps o 3 satellites: 2 x 832bps = 1664bps Decentralised in each satellite (no information about the local satellite coordinates needs to be sent, since it is the local reference frame (GFF)): o 2 satellites: 309bps o 3 satellites: 2 x 309bps = 618bps Table A-7 in appendix A provides the RF metrology data description of the RF IF to the GNC subsystem, illustrated only for 1 satellite (see RD[8]). RF metrology data from the GNC subsystem: centralised in the Coronagraph and decentralised in each satellite: o 3 satellites: 3 x 204bps = 612bps o 2 satellites: 2 x 204 bps = 408bps IF between generated RF navigation data and TT&C The generated RF navigation data has to be transferred to Earth for experimentation. Table A-8 in appendix A shows the data for the case of two and three satellites (see RD[8]). The RF metrology interacts with the TT&C subsystem in order to transfer the generated data to Earth for on ground experimentation, which includes the generated raw data and the PVAT navigation, at 1Hz The communication between the RF metrology and the TT&C subsystem is done via the OBDH at 1Hz. Both subsystems are located in the same satellite. The total RF navigation data that needs to be transferred to Earth via the TT&C subsystem for experimentation is: 3 satellites: 856bps 2 satellites: 570bps Conclusion: IF data between RF metrology and, OBDH and TT&C The mission with two satellites: Decentralised mode: o Total data to be processed by the OBDH = (RF to GNC) + (GNC to RF) + (RF to TT&C) = = 1287bps

48 Page 46 of 231 o Total data to be sent to Earth via the TT&C in the return link: (RF to TT&C) = 570 bps Assuming the mission with three satellites: For deployment in decentralised mode: o Total data to be processed by the OBDH = (RF to GNC) + (GNC to RF) + (RF to TT&C) = = 2086bps o Total data to be sent to Earth via the TT&C in the return link: (RF to TT&C) = 856bps For nominal operation in centralised mode: o Total data to be processed by the OBDH = (RF to GNC) + (GNC to RF) + (RF to TT&C) = = 3132bps o Total data to be sent to Earth via the TT&C in the return link: (RF to TT&C) = 856bps RF metrology antenna trade-off The requirements for the RF S-band antennas are: Same antenna is to be used for both carrier frequencies and for Tx and Rx Phase centre accuracy < 1mm (rms) for nominal, and < 5mm (rms) for deployment Antenna gain (omni-directional) is to be higher than 3dB over the off-axis angle [-90 ;+90 ], in order to allow full sphere (4π steradian) visibility. Three types of antennas have been considered (see RD[7]). Kind of antenna Size Mass Gain at bore-sight Gain at 90 Cross dipole Quadrifilar helix antenna Patch antenna D=10cm H=4cm D=10cm H=20cm D=8cm H=1cm 0.2 Kg 4 dbi -5 dbi 0.3 Kg 3 dbi -3 dbi 0.1 Kg 4 dbi -5 dbi Table 6-4: Trade-off for antenna selection, typical performances Quadri-helix antennas are recommended because they: Have superior field of view (omni-directional). Full visibility with 2 sets of antennas (-3dB assumed in the link budgets) Provide good circular polarization Have a bandwidth larger than 200MHz for dual frequency (110MHz separation between the two frequencies S1 and S2) Offer the possibility to calibrate the phase centre down to a few mm GPS requirements and manufacturers selection GPS single frequency (L1) is needed on board all the satellites for three reasons: Absolute GPS, in order to synchronise the OBDH subsystem of the satellites.

49 Page 47 of 231 Absolute GPS, to provide absolute positioning for each satellite in the formation, in order to aid the RF metrology acquisition in case of problems and to serve as backup for the anti-collision functionality (10 meters accuracy would be sufficient) Relative GPS (carrier phase) post-processing on ground (cm level accuracy), for FF RF metrology performance validation for ISD beyond 1 km (no optical link available) Typical R-GPS performances are presented in Figure 6-1 (RD[9]). Typical absolute GPS performances are: Position: 10-20m 1-σ (3D-Pos) Time: 1µs 1-σ Because of the orbit selected for, the use of GPS is only possible during the perigee passage (below 5000 Km); therefore a navigation propagator has to be implemented to cover these visibility holes. Two antennas are needed, mounted on the +Z and Z sides, to allow continuous GPS satellite acquisition and tracking and to avoid occultation due to the other satellites in the formation. The GPS raw data (measurements) will be sent to Earth via the TT&C link to be processed, estimated to be around 200bps per satellite. Figure 6-1: Typical R-GPS performances Several companies in Europe provide single frequency GPS receivers. Table 6-5 provides a list of these companies and GPS products with the main characteristics.

50 Page 48 of 231 Company Product Main characteristics Alcatel TopStart 3000 LEO, re-entry, GTO, GEO Space PVT, attitude measurements, High sensitivity tracking L1(now) & L1/L2C(2006) Laben GPS Tensor LEO PVT, attitude L1 Lagrange LEO Precise positioning, scientific L1/L2 Surrey Satellite Tecnolgy Ltd (SSTL) SAAB / Austrtian Aeorspace (AAE) Astrium Germany German Space Center (DLR) SGR-05, -10, -20 LEO Low cost PVT, attitude (SGR-20) L1, low power, minimum size/mass, based on COTS GRAS LEO PVT, Atmospheric sounding receiver L1/L2 InnovativeGPS (2007) LEO, GEO Precise PVT, attitude, low power and mass. L1/L2C Mosaic GNSS LEO, GTO, GEO PVT, attitude, highly sensitive, L1, high integration in on board avionics Phoenix LEO, re-entry, launchers Low cost PVT, range safety functions, KGPS of FF L1, low power, minimum size/mass, based on COTS Table 6-5: European companies with GPS products, main characteristics Two products are selected due to the low cost, low power consumption and low size/mass, the SGR-05 from SSTL (2 RF antennas inputs) and the Phoenix receiver from DLR (option due to the need of an external RF combiner to allow the use of two external RF antennas). 6.3 Two Satellites Solution (Baseline) The RF metrology equipment on the satellites is exactly the same for the two satellites, the Coronagraph and the Occulter. The equipment is described only once for both satellites RF metrology subsystem unit description The FF RF metrology subsystem unit for each satellite consists of: 2 sets of 3 RF antennas, an Rx/Tx unit based on GPS receiver heritage (TopStar 3000 from Alcatel space), and the Navigation Processing Unit (NPU) embedded in the Rx/Tx unit. The FF RF metrology unit is being developed by Alcatel Space with GMV, based on Darwin and SMART 2/3 (RD[6]) requirements, and it will be used in the PRISMA mission. The Rx/Tx unit broadcasts FF RF signals (for use by other Rx/Tx units), receive FF RF signals (from others Rx/Tx units) and provide measurements to the NPU. The NPU computes ranging and angular measurements (elevation and azimuth) within the formation, and passes this information to the GNC algorithms. Figure 6-2 illustrates the expected final view of the EM for each RF metrology unit, with the Tx/Rx and NPU integrated in a single mechanical box.

51 Page 49 of 231 Figure 6-2: View of the BB01 unit used during BB activities (left) with the Tx board stacked on the TopStar3000 EM (Alcatel Space). The final RF metrology EM will have a similar look, with the Tx board (right) and the NPU SW embedded in one single unit Measurements consist of pseudoranges at two frequencies (P1 and P2) between transmitters and receivers of different satellites, and single difference between carrier phases (S1 and S2) for the multiple antennas located on the same satellite. These measurements allow the NPU to compute the relative navigation within the formation, passing the PVAT navigation solution in Cartesian and polar coordinates, to the GNC subsystem. FF RF metrology subsystem operation, and consequently the NPU processing, is distributed in deployment mode and centralized in nominal FF mode. The Rx and Tx sub-units of a single satellite share a common clock, in order to have a unique clock bias. The RF metrology unit is composed of (see Figure 6-3): receiving RF hardware, digital baseband chip, microprocessor and associated memory, transmitting up-converters and RF power amplifiers, OCXO and DC/DC converter. The equipment is equipped with 6 RF S- band antennas, comprising 2 Rx/Tx antennas and 4 Rx-only antennas. The baseline for the antenna type is S-band quadri-filar helix. Figure 6-3: RF metrology architecture for, with 2 sets of 3 S-band antennas

52 Page 50 of 231 The expected budgets for the final EM are (RD[6]): RF metrology unit budget (one per satellite): o mass 5.4Kg o size 290*180*130mm o power consumption 29W (deployment), 20W (nominal) o electrical IF 5VDC, 6VDC and 15VDC (TBC) o RS422 serial links RF S-band antennas, 6 antennas (2 Tx/Rx + 4 Rx). Budget per antenna: o Length approx. 70mm without connector o Reflector d = 100mm o Radome d = 60mm o Weight 160g RF metrology navigation performances The navigation performances presented in Table 6-6 are based on the FF RF metrology BB results for a formation of two satellites (equivalent to the mission baseline) (RD[6]). Performance degrades for increasing azimuth and elevation angles. The antenna baseline for the angular measurements was assumed to be 1m with the BB test activities. A degradation factor of 20% in the angular measurements accuracy performances has to be included in the mission due to the dimensions of the satellites (80x90x90cm). The performance of the RF metrology guarantees the acquisition of the optical metrology, that requires 1cm longitudinal accuracy and +/-1 lateral accuracy. Position accuracy (99.9%) 2-satellite Range 0.7 cm Azimuth 4.1 (line-sight <30) 1 (line-sight >30) Elevation 3D (r=2km, line-sight <30) Velocity accuracy (99.9%) 1.1 (El<60 ) 2.5 (El<80 ) <44 m Range rate 1.4 mm/s Azimuth rate 0.05 /s (line-sight <30) 0.09 /s (line-sight >30) Elevation rate 0.08 /s (El<60 ) 0.2 /s (El<80 ) 3D (r=250 m) - Euler angles accuracy (99.9%) in 321 sequence Attitude (r=250 m) - Attitude rate -

53 Page 51 of 231 Time To Fix Formation Navigation Worst case: quasi-static conditions TTFF: first fix TTFA: with nominal performances 2-Nodes Initialisation time TTFF (95%) warm start TTFF (95%) cold start TTFA (95%) warm start TTFA (95%) cold start < 60 s <51 s <300 s < 151 s < 400 s Table 6-6: Navigation accuracy (99.9%) measured with the FF RF metrology BB, for the case of two satellites (nominal formation). Maximum range with a fixed TM/TC data rate of: Master (Coronagraph): 9kbps; Free Flyer (Occulter): 3kbps (Table 6-7). Table 6-7: Maximum range improvement with fixed TM/TC IFs between RF metrology and, OBDH and TT&C subsystem For the mission with two satellites, working in decentralised mode: Total data to be processed by the OBDH = 1287bps Total data to be sent to the Earth via the TT&C in the return link = 570bps RF metrology S-band antennas The recommended antennas, due to their low mass and small dimensions, are the S-band antennas used for communications in Proba-2, provided by STT. See RD[5] for detail specifications.

54 Page 52 of 231 Figure 6-4: STT S band antenna used in Proba-2 (left) and radiation patter (right). Main performances of this antenna are (over the operating temperature range -70 C to +60 C): Electrical: o Frequency Range MHz uplink o MHz downlink o Gain 3dBic ± 1dB boresight, typ. -6dBic ±2dB for 90 < Θ < +90, typically o Polarization Right circular (RHC) Mechanical: o Length approx. 70mm without connector, o Reflector d = 100mm, o Radome d = 60mm; o Weight 160g RF metrology open points and critical issues The RF Metrology subsystem is a major technology step forward to be demonstrated in. The RF final performances cannot be assessed on ground, mainly because of the difficulty of reproducing the operational conditions in space (relative dynamics, multipath, RF environment). Several critical issues have been identified during the BB test activities (RD[6]): A precise calibration and stability of position of the antenna phase centre is needed (1mm level). This is a matter to be demonstrated through measurements in an anechoic chamber with a satellite model, but it will have to be further validated in space. A precise calibration and stability of Rx/Tx hardware biases are needed (1mm level). Feasibility was demonstrated during phase 2 of the BB activity, but it will have to be further validated in space. FF RF subsystem ground tests are limited by poor sensor characterisation due to multipath effects under space environmental conditions. Flight demonstration is therefore mandatory. Final performance will be highly sensitive to multipath mitigation performance. It is of course expected that the satellite design will be constrained by other drivers than multipath mitigation and therefore will not be optimal to this respect (for example number of edges, reflection coefficient of surface). It has been shown that to get reasonable performance, about 70% mitigation is needed at software level, for carrier phase measurements. The only method deemed realistic is based on repeatability of

55 Page 53 of 231 multipath. It leads to the calibration of the multipath error from ground measurements performed in an anechoic chamber GPS unit description GPS single frequency (L1) is used on board all the satellites for synchronisation of the OBDH subsystem of the satellites, absolute positioning for each satellite, and relative GPS (carrier phase) post-processing on ground (cm level accuracy), for FF RF metrology performance validation. The GPS receiver baseline is the SGR-05 from SSTL (2 RF antennas inputs). The performances are: navigation o 3D-Pos: 10-20m, 1-σ o Time (UTC): 1µs, 1-σ o TTFF: 60 s Receiver budgets: o Volume : 76x246x170mm o Mass: 300g (including mechanical box) o Nominal Power consumption: 2 W Antennas (2 for full coverage), per unit: o Mass: 200g o Dimensions: 100mm, height 20mm Figure 6-5: SSTL SGR-05 OEM board (left), and DLR phoenix board (right) Recommended requirements for the RF metrology subsystem demonstration Functional & operational requirements for demonstration The functionalities of the RF FF metrology subsystem to be demonstrated in space are: Computation of relative navigation During the deployment, two basic functions, collision avoidance and coarse position

56 Page 54 of 231 Provision of a synchronized clock reference to each satellite Local inter-satellite data link (TM/TC) Autonomy and robustness. Full sphere (4π steradian) visibility Transition between RF metrology and optical metrology Manoeuvre requirements for the RF metrology demonstration Deployment at r (sphere radius) = 15km in the worst case of power link budget (only ranging function without demodulation) and with the worst geometry from the FF RF metrology subsystem point of view (90º rotation around one axis) Deployment at r (sphere radius) = 8km (maximum range for deployment) in the worst case geometry from the FF RF metrology subsystem point of view (90 º rotation around one axis) Demonstration of full coverage capability by rotation of 360º around two axes, at 8km ISD Performance validation during deployment from 8km (maximum range for deployment) down to 250m (maximum range for nominal formation), with rotations of the satellites < 5deg/s, around two axis. The distance between satellites is progressively decreased in several steps as required Performance characterisation during nominal manoeuvres, from 250 meters (maximum range for nominal formation) down to 5 meters (minimum distance between satellites for collision avoidance), and under different manoeuvres modes: Coarse Formation and Reconfiguration Modes Performance characterisation during nominal mode, in quasi-static conditions, from 250m down to 5m, in Baseline Mode. Initialisation of the FF RF metrology, relative positioning fixed in the formation in nominal mode, in quasi-static conditions, at 250m and at 25m, in Baseline Mode RF near-far field demonstration. Two satellites located at 25m, and the third satellite located at 8 km from the other two. Initialisation and relative positioning fixed within the time requirements mission requirements in support of the RF metrology demonstration The following elements are required to be embarked on the satellites in order to carry out the inflight technology validation. Availability of the optical metrology system with mm accuracy, for ISDs between 1km and 5m, for the FF RF performance characterisation Availability of GPS receivers on board the satellites, for the R-GPS (carrier phase) postprocessing on ground (cm level accuracy), for FF RF performance validation for ISDs over 1km Availability of a ground-space satellite link, to be used for experimental data collection and transmission to Earth Hybridization of the FF RF subsystem with other attitude sensors (for instance, star trackers) for the case of two satellites.

57 Page 55 of List of Equipment The RF metrology main elements are: RF metrology unit: Alcatel Space and GMV. It is used in PRISMA, and BB developed at ESA under contract n 15511/02/NL/EC (RD[6]). See BB in Figure 6-2 S-band quadric-helix antennas: SST provider for Proba-2. See RD[5] and Figure 6-4 GPS receiver: SGR-05 from SSTL provided with 2 RF antennas (see RD[11]), see Figure 6-5 GPS antennas: GPS S Series. Sensor Systems Inc. USA. See RD[12] and RD[13]. 6.5 Configuration The S-band RF antennas are located in the two planes perpendicular to the +X axis (3 antennas in each opposite side of the satellite), without RF obstructions, allowing full visibility with two sets of three antennas (-3dB assumed in the link budgets). The disposition of each set of three antennas is forming a triangle, with maximum distance between the antennas. This geometry optimises the navigation accuracy performances. Optical sensors are placed in the same planes, where the performance is optimised (performance improves with larger antenna baselines, regular geometries (equilateral triangle) and with number of satellites (three is better than two)). Hybridisation with attitude sensors (star tracker) is needed for the case of two satellites (during deployment) to know the relative attitude. The two GPS L1 patch antennas are mounted on the +Z and Z sides, to allow continuous GPS satellite acquisition and tracking and to avoid occultation by the other satellites in the formation. The GPS raw data (measurements), 200bps per satellite, is sent to Earth via the TT&C link Mass& dimensions budgets Element 1 Occulter MASS [kg] Unit Unit Name Click on button above to insert new unit Quantity Mass per quantity excl. margin Maturity Level Margin Total Mass incl. margin 1 RF Formation Flying subsystem To be developed S-band LGA (for FF) Fully developed GPS receiver Fully developed GPS antenna Fully developed ISL transceiver To be developed ISL antennas To be modified TM/TC unit (S-band transceivers) Fully developed S-band LGA (for TT&C) Fully developed To be developed SUBSYSTEM TOTAL Table 6-8: Total mass budget of the communications subsystem, where RF metrology is included

58 Page 56 of 231 Element 1 Occulter DIMENSIONS [m] Unit Unit Name Quantity Dim1 Dim2 Dim3 Click on button above to insert new unit Length Width Height or D 1 RF Formation Flying subsystem S-band LGA (for FF) GPS receiver GPS antenna ISL transceiver ISL antennas TM/TC unit (S-band transceivers) S-band LGA (for TT&C) SUBSYSTEM TOTAL 6 Table 6-9 Dimensions of the communications subsystem, where RF metrology is included 6.6 Options RF Metrology for the three satellites option The FF RF metrology on board equipment description is the same for two or three satellites at HW level, and the information provided for the two satellite baseline is applicable. The only difference is at SW level concerning the navigation processing (algorithms), that is significantly different for a two (1D) or three (2D) satellites formation. The use of a third satellite for would lead to some modifications in terms of performances and data rate Third satellite added value The RF metrology final performances cannot be assessed on the ground, mainly because of the difficulty in reproducing the operating conditions in space (relative dynamics, multipath, RF environment). The use of a third satellite would provide the following added value: RF multipath effects cannot be accurately analysed on-ground and require three (or more) satellites in space to be fully representative. Nevertheless, RF multipath characterisation with 2 satellites can give information to update the RF models used onground. For the operational and performance aspects of the RF metrology system, the use of a third satellite is very valuable. The navigation algorithms of the RF metrology sensor are significantly different for a two or three satellites formation, and the final performance depend on the RF navigation algorithms and the number of measurements available for the navigation filter RF metrology navigation performances The navigation performances presented in Table 6-10 are based on the FF RF metrology BB results for a three satellites formation ( mission option) (RD[6]). The addition of a third satellite allows for maintenance of planarity of the formation, so that relative azimuth and elevation angles always remain within acceptable levels. The case of a FF mission with four satellites is just provided for information. The antenna baseline for the angular measurements was assumed to be 1m with the BB test activities. A degradation factor of 20% in the angular

59 Page 57 of 231 measurements accuracy performances has to be included in the mission due to the dimensions of the satellites (80x90x90cm). Position accuracy (99.9%) 3- SATELLITE 4- SATELLITE Range 0.7 cm 0.2 cm Azimuth Elevation 3D (r=2km, line-sight <30) 1.1 (El<60 ) 2.5 (El<80 ) 0.05 <44 m 2.3 m Velocity accuracy (99.9%) Range rate 1.4 mm/s 0.4 mm/s Azimuth rate 0.05 /s /s Elevation rate 0.08 /s (El<60 ) 0.2 /s (El<80 ) 0.03 /s 3D (r=250 m) 20.5 cm/s 3.0 cm/s Euler angles accuracy (99.9%) in 321 sequence Attitude (r=250 m) /0.11 /0.23 Attitude rate - 30 mdeg/s Time To Fix Formation Navigation Worst case: quasi-static conditions TTFF: first fix TTFA: with nominal performances Initialisation time TTFF (95%) warm start TTFF (95%) cold start TTFA (95%) warm start TTFA (95%) cold start 3/4-Nodes < 60 s <281 s <530 s < 1921 s < 2170 s Table 6-10: Navigation accuracy (99.9%) measured with the FF RF metrology BB, for the case of three satellites (option). Maximum range with a fixed TM/TC data rate of: Master: 9kbps; Free Flyer: 3kbps (Table 6-11). Table 6-11: Maximum range improvement with fixed TM/TC (3 satellites) IFs between RF metrology and, OBDH and TT&C S/S For a mission with 3 satellites, working during deployment in decentralised mode and during nominal formation in centralised mode: For deployment: o Total data to be processed by the OBDH = 2086bps

60 Page 58 of 231 o Total data to be sent to Earth via the TT&C in the return link = 856bps For nominal: o Total data to be processed by the OBDH = 3132bps o Total data to be sent to Earth via the TT&C in the return link = 856bps GPS optional manufacturer The GPS receiver option is the Phoenix receiver from DLR (to fly on PROBA2). See RD[13] and Figure 6-5.

61 Page 59 of OPTICAL METROLOGY 7.1 Requirements and Design Drivers The optical metrology system is a highly accurate satellite position (overlay) sensor, which is used by the GNC system of the satellite formation to improve its FF accuracy beyond what is achievable with sensors based upon RF metrology. However, the optical metrology system cannot operate without an initial formation alignment as provided by the RF metrology. The optical metrology system in its present form is not designed to measure relative satellite attitude information; it is assumed instead that commercial star trackers are used for this purpose, which can provide absolute attitude information down to a few arcsec accuracy. The optical metrology system was initially developed and breadboarded for the DARWIN formation of 7 satellites (TRP activity 15645), but it is very well applicable to the requirements. It consists of the following sub-systems: Coarse lateral sensor Fine lateral sensor Longitudinal sensor Fringe tracking sensor 7.2 Assumptions and Trade-Offs As mentioned in the previous Section, the optical metrology system can only be used if the alignment of the satellite formation is already controlled to the following accuracies between the Coronagraph satellite and any other satellite (for example the Occulter satellite): Attitude or positional error Roll, pitch, yaw Lateral (positional offset) Longitudinal (distance offset) Value < ±10 arcsec <±1.5 degrees <±25 mm Table 7-1: Accuracy of the RF metrology

62 Page 60 of 231 Figure 7-1: Satellite alignment with RF metrology The inter-satellite distance is assumed to be no more than 250 meters. The angular and positional alignment between two satellites, with RF metrology sensors is indicated in Figure Coarse lateral sensor (CLS) The CLS enables the lateral misalignment angle between the Coronagraph and the Occulter satellites to be reduced to 1 arcsec, which corresponds to 0.5 mm at 100 meters ISD. The CLS is based on a diverging laser beam, which is retro-reflected from a corner-cube and the position of the retro-reflected light measured by a CCD camera with tele-objective lens. Its principle design is shown in Figure 7-2: Figure 7-2: CLS principle

63 Page 61 of 231 The technology of the CLS is well established and therefore no breadboarding was performed in the frame of the HPOM TRP activity. The angular and positional alignment between two satellites, with RF and CLS is indicated in Figure 7-3. Figure 7-3: Satellite alignment with RF and CLS Fine lateral sensor (FLS) The FLS is an optional sensor that can only operate once the CLS has increased the lateral accuracy initially provided by the RF metrology. The FLS enables the lateral misalignment between the Coronagraph and the Occulter satellite to be further reduced, down to ±35 micrometers. A principle scheme of the FLS is shown in Figure 7-4. It uses a collimated beam transmitter and the CCD or PSD camera to sense the lateral offset. This is a technology, which does not scale with the ISD. Figure 7-4: FLS design principle

64 Page 62 of 231 An image of the breadboard is shown in Figure 7-5, which shows the whole setup on the right and the laser transmitter in detail on the left. Figure 7-5: FLS emitter (left) and full FLS system (right) shown together with the DWI on an optical bench The angular and positional alignment between two satellites, with RF, CLS and FLS is indicated in Figure 7-6. Figure 7-6: Satellite alignment with RF, CLS and FLS Absolute longitudinal sensor The absolute longitudinal sensor is based on interferometry. The interference phase in any interferometer is proportional to the optical path difference (OPD) in both of the interferometer arms and inversely proportional to the laser wavelength. Thus, if the OPD is increased from zero

65 Page 63 of 231 until it corresponds to the wavelength, the interference phase increases from 0 to 360 degrees, after which the interference phase and thus the target distance become ambiguous. As laser wavelengths are very short, so is the ambiguity range. To overcome the problem two wavelengths are used. In a dual-wavelength interferometer (DWI) the beat (difference) frequency of two ultra-stable lasers is measured by a high-speed photodiode, compared to a stable oscillator and one of the two lasers is servo controlled to maintain the frequency difference. An interferometer is fed with both laser wavelengths, which renders the phase difference between interferences of both wavelengths. This phase difference is still proportional to the OPD, but inversely proportional to a synthetic wavelength. The synthetic wavelength is given by the product of the two individual laser wavelengths and divided by their difference. Thus, if the laser wavelengths are close, the synthetic wavelength can be made large. In the case of the DWI built for HPOM, the wavelength difference is so small that it is better expressed in a frequency difference, namely 3 GHz. This in turn corresponds to a synthetic wavelength of 100 mm and this is the un-ambiguity range of the DWI. It is interesting to note that the stability of the individual laser frequency is irrelevant in a DWI; it is only the (electrically controlled) difference in frequency that determines the stability of the interferometric measurements. In order to perform accurate DWI phase measurements heterodyning was implemented. Four acousto-optical modulators (two per wavelength) create heterodyne beat frequencies of 1 MHz (for laser 1) and 1.5 MHz (for laser 2). The heterodyne frequency generation scheme and a picture of it are shown in Figure 7-7. Figure 7-7: DWI heterodyne frequency generation scheme and picture (please ignore the optical components in the upper left of the picture - on the separate small optical bench -, they do not belong to the frequency generation equipment) Two polarization maintaining single-mode fibres carry the heterodyne frequency shifted laser beams to the optical head of the DWI. Both detectors in the optical head are subjected to the same laser and heterodyne frequencies. Their only difference is the interferometer phase, which is introduced by the longitudinal distance to be measured. Figure 7-8 shows the principle layout of the interferometric optical head on the left and a photograph of the breadboard on the right.

66 Page 64 of 231 Figure 7-8: DWI head, principle layout and picture. The delay line shown in the layout is not shown in the picture. Please note that the corner-cube at the end of the distance to be measured is very close to the interferometer head for testing. The angular and positional alignment between the two satellites, with RF, CLS, FLS and DWI is indicated in Figure 7-9. Please note that the FLS is not required for the operation of the DWI. Figure 7-9: Satellite alignment with RF, CLS, FLS and DWI

67 Page 65 of Fringe tracking sensor (FTS) One of the two lasers of the DWI has been chosen to give a frequency-doubled output (a wavelength of 532 nm), which is used to stabilise the laser to an iodine absorption line. This Doppler-free Pound-Drewer wavelength stabilization technique enables stabilities in the order of 10-12, which is necessary to achieve a distance resolution of 5 nm over distances of 500 meters. Figure 7-10 shows the schematics and a picture of the Pound-Drewer stabilization breadboard. Figure 7-10: Pound-Drewer frequency stabilization method for Nd:YAG laser sources The angular and positional alignment between two satellites, with RF, CLS, FLS and FTS is indicated in Figure Please note that the FLS is not required for the operation of the FTS, nor is the DWI. Figure 7-11: Satellite alignment with RF, CLS, FLS and FTS

68 Page 66 of Two Satellites Solution (Baseline) Figure 7-12 shows the distribution of the optical metrology sensors between two satellites, with the following symbols: White box: Corner-Cube Retro-reflector (CCR) Red box: Coarse Lateral Sensor (CLS) Blue box: Absolute Longitudinal Sensor (DWI) and Fringe Tracking Sensor (FTS) Brown box: Fine Lateral Sensor (FLS) transmitter Black box: Fine Lateral Sensor (FLS) receiver Please note that the CCR is utilised simultaneously by the CLS, the DWI and the FTS. As the systems use different laser wavelengths, a dichroic beam splitter is indicated on the Occulter satellite to superimpose the beams. Figure 7-12: Optical metrology system distribution in the two satellites solution The optical metrology equipment is distributed between the two satellites with the following optimisations in mind: 1. Minimisation of the metrology data flow between the satellites 2. Reduction of straylight into the Coronagraph satellite. The optical metrology performance parameters and the volume, mass and power figures are given in Table 7-2 and Table 7-3 respectively. Metrology instrument Measured/Expected performance Remarks Coarse lateral sensor (CLS) 1 ±1 degree, 10 Hz Fine lateral sensor (FLS) 10 µm/ ±5 mm, 10 Hz Absolute longitudinal sensor (ALS) 10 µm/ 250 m, 10 Hz * Fringe tracking sensor (FTS) 2 nm/ 250 m, 10 Hz ** Remarks: * Makes additional use of RF metrology system to obtain absolute distance ** Iodine line stabilisation implemented in DWI Table 7-2: Measured (or expected) performance of the optical metrology sensors

69 Page 67 of 231 Metrology instrument Volume [mm] Mass Power Remarks Corner cube retro-reflector (CCR) Diameter 30 x kg 0 Watts * Coarse lateral sensor (CLS) 200 x 100 x kg 8 Watts Fine lateral sensor (FLS) transmitter 50 x 50 x kg 3 Watts Fine lateral sensor (FLS) receiver 50 x 50 x kg 2 Watts Absolute longitudinal sensor (ALS) 400 x 200 x kg 27 Watts Fringe tracking sensor (FTS) 300 x 200 x kg 27 Watts Both ALS and FTS combined 400 x 300 x kg 35 Watts ** Remarks: * The CCR is used by the CLS, the DWI and the FTS simultaneously ** Considerable mass and power savings can be achieved if the DWI and the FTS are operating simultaneously, because equipment (optical head, laser and so on) can be reused. Table 7-3: Interface requirements of the optical metrology equipment Coronagraph satellite The Coronagraph satellite is carrying: 1. CCR 2. FLS transmitter The total optical metrology volume, mass and power data of the Coronagraph satellite is given in Table 7-4. Metrology instrument Volume [mm] Mass Power Corner cube retro-reflector (CCR) Diameter 30 x kg 0 Watts Fine lateral sensor (FLS) transmitter 50 x 50 x kg 3 Watts Total: 0.6 kg 3 Watts Table 7-4: Coronagraph satellite optical metrology package requirements Occulter satellite The Occulter satellite is carrying: 1. CLS 2. FLS receiver 3. DWI 4. FTS The total optical metrology volume, mass and power data of the Occulter satellite is given in Table 7-5. Metrology instrument Volume [mm] Mass Power Coarse lateral sensor (CLS) 200 x 100 x kg 8 Watts Fine lateral sensor (FLS) receiver 50 x 50 x kg 2 Watts Absolute longitudinal sensor (DWI) 400 x 200 x kg 27 Watts Fringe tracking sensor (FTS) 300 x 200 x kg 27 Watts DWI and FTS combined 400 x 300 x kg 35 Watts Total with DWI or FTS 22.2 kg 37 Watts Total with DWI and FTS 32.2 kg 45 Watts Table 7-5: Occulter satellite optical metrology package requirements

70 Page 68 of Options Only the CLS is required to perform the coronagraph experiments with two satellites. All other optical metrology sensors are optional and intended to be technology demonstrations. The FLS and the DWI would demonstrate the formation s ability to maintain a FF accuracy (in three dimensions), which is required for the DARWIN mission. The FTS would improve the longitudinal (distance) resolution by an additional 4 orders of magnitude and clarify whether FF down to nanometre stabilities is feasible. This would answer the question whether delay lines will be required for DARWIN. 7.4 Three Satellites Option There are two optical metrology options with three satellites. One where the optical metrology is implemented between the Coronagraph and each one of the two free-flyers as shown in Figure 7-13 on the left, and a full fledged version, where additional optical metrology is implemented between the two free-flyers for redundancy and accuracy enhancement as shown in Figure 7-13 on the right. Figure 7-13: Optical metrology system distribution in the three satellites solution

71 Page 69 of SUN OBSERVATION PAYLOAD (CORONAGRAPH) 8.1 Introduction This Chapter is meant to provide the reader with a short overview of the Sun observation payload, considered in the frame of the Study and of its requirements and impact on the mission and spacecraft design. Its content is based on the information and inputs provided by Laboratoire d'astrophysique de Marseille. More detailed information related to the proposed instrument is contained in RD[17] 8.2 Coronagraph Science Objectives The principal objective of Solar corona observation, using an optical system distributed on two satellites, can be summarized as: To characterize the physical processes in the inner corona, as: o Acceleration of the solar wind (2D velocity map) o Propagation of Corona Mass Ejections (CMEs) o Wave propagation o Turbulence o Magnetic field. The performance of a coronagraph is mainly driven by the distance between the external occulter and the entrance pupil. The further away the external occulter is from the entrance pupil, the less stray light will affect the measurements. This characteristic makes the use of a FF mission an interesting tool for a coronagraph instrument, providing the possibility of flying the external occulter of the instrument on a separate satellite. The operating principle of a coronagraph is shown in Figure 8-1 below. Figure 8-1: Principle of a coronagraph. For more information regarding the Solar corona observation and the coronagraph considered for, refer to RD[17]. 8.3 System Requirements and Design Drivers The main system requirements and design drivers coming from the coronagraph instrument are given by: The relation between the occulter disc and the ISD, having a value of 1/100. For example, a distance between the two satellites of 120m yields to an external occulter disc diameter of 1.2m

72 Page 70 of 231 The absolute stability and the pointing accuracy of the Coronagraph and the Occulter, values are given in Table 8-1 The lateral and longitudinal relative position between the two satellites. The requirements given in Table 8-2 allow observations of the Corona between 1.05RSun and 3.2RSun, RD[18] The exposure time, always lower than 15s, being the nominal exposure time for the monochromatic images equal to 10s and for the polychromatic images 1s. The thermal requirements of the telescope optics, which should be kept at around 20ºC, and of the CCD detector, to be maintained at -70ºC, while keeping the temperature gradient of the optical bench as low as possible. The amount of data produced by the instrument. Each composite image is formed by 1 polychromatic image and 4 monochromatic ones, having the following sizes: o Polychromatic: 6.7Mbits/image (compression 10), o Monochromatic: 2.2Mbits/image (compression 30). Adding up to a total image size of 15.5 Mbits/image. On each coronagraph science dedicated orbit, observation of the Sun will be carried out over a 12 hours long orbital segment centred across the orbit s apogee, with an instrument s maximum on-time envisaged to be equal to six hours within this 12 hours period. Furthermore, for sizing, it has been considered that the instrument will take a full set of pictures every five minutes. Pointing Requirements Absolute stability Pointing accuracy Coronagraph S/C Occulter S/C Polychromatic 1 [arcsec/image] (1 image = 1s) No such requirement Monochromatic 5 [arcsec/image] No such requirement (1 image = 10s) Pitch/Yaw ±40 [arcsec] <2 [º] Roll ±0.5 [º] No such requirement Table 8-1: Pointing requirements for the Coronagraph and Occulter S/Cs. Positioning Requirements Lateral [mm] ±2.5 Longitudinal [mm] ±250 Table 8-2: Positioning requirements for the Formation.

73 Page 71 of Coronagraph Design Overview At overall system level, the coronagraph will consist of elements on both the satellites of the formation. The only element embarked on the Occulter satellite will be the external occulter disk, whose diameter, linked to the ISD distance, is envisaged to be equal to 1.2 meters. The instrument items embarked on the Coronagraph satellite will consist of two modules, the telescope and an electronics box. The overall mass and power requirements of this composite, together with the dimensions of these two boxes are given in Table 8-3. Physical parameters Instrument Mass [kg] 30.1 (incl.20% margin) Power [W] 20 (without thermal control) Telescope Dimension 1 [mm] 900 Dimension 2 [mm] 300 Dimension 3 [mm] 300 Electronics box Dimension 1 [mm] 200 Dimension 2 [mm] 200 Dimension 3 [mm] 200 Table 8-3: Physical parameters of the Coronagraph instrument

74 Page 72 of 231 This Page Intentionally Blank

75 Page 73 of GUIDANCE, NAVIGATION and CONTROL This chapter will define the on board Guidance, Navigation and Control (GNC) system proposed for the satellites. It will define the modes, the relative dynamics and the propulsion selected. After this follows the definition of manoeuvres and their V computations and then the trade off between two or three satellites for Formation Flying, finishing with the definition of the sensory avionics equipment. 9.1 GNC Main Modes The system modes described earlier are seen and driven by the different implemented subsystems on board. The modes presented here are the functional modes seen from and needed in the GNC system. The state diagram is illustrated in Figure 9-1. It shall be noted that the GNC system needs to work in both a centralised and a decentralised top level mode. In the centralised mode the GNC system runs on one satellite and exchanges measurements, commanded forces and torques with the other. The formation is considered as one complete system. In the decentralised mode, each satellite has a full GNC system, which exchanges information with the other one, but shall also be able to operate autonomously in closed loop without the other satellite. The formation is considered as two separate systems communicating. What is allocated to each mode is illustrated in Figure 9-1. Decentralised mode Centralised mode Figure 9-1: Main GNC modes state transition diagram. The modes in Figure 9-1 are the following: Separation: This mode is activated while the satellites are still attached to the launch structure and the on board systems, including the RF, is activated and initialized. The satellites are then jettisoned.

76 Page 74 of 231 Formation Flying Deployment: After the separation, the satellite detumbles and brakes to Station Keeping (SK) position. The former satellite might have to relocate with respect to the expected position of the latter to avoid collision risk. When all satellites are at rest they will autonomously be brought together to form the default rest formation. Safe Mode: All anomalies which affect or endanger the controllability of any satellite shall trigger the safety system, which will autonomously bring the formation in a safe mode. This will mean nulling all relative velocities and if not possible, the healthy satellite will move out of the trajectory of the unhealthy one. Operations: A standby state, where the satellite is able and ready to perform all other states described in Figure 9-1. Perigee Pass: This part handles the part of the orbit which is closer to Earth and has higher disturbances. The formation is dissolved and the passage from one side to the other is performed as the standard Rendezvous 2-pulse hopping manoeuvre. This manoeuvre is stopped at the end with the second pulse. There is closed loop control along the 2-pulse manoeuvre trajectory used for guidance. If it is left open loop, the co-variance ellipsoid for the dispersion at the final point will be too large, with respect to the nominal target location. Formation Flying Manoeuvres: This mode will perform all the various manoeuvres to be tested for the Formation Flying. It will in practice consist of many submodes not illustrated in Figure 9-1, but which is pertinent to the GNC architecture. The manoeuvres are described in Chapter Coronagraph Science Observations: This mode handles all operations of the Coronagraph with respect to GNC. The formation is basically a very simple inertial SK manoeuvre at a fixed distance between the two satellites and does not require any manoeuvring at all. This mode will occupy the majority of the total mission duration. 9.2 Orbital Dynamics This section will briefly describe the orbital parameters needed for the GNC, the frames used for the dynamics and its points of concern as well as the propulsion issues Orbital data The orbit is highly eccentric with a period of 24h and a perigee altitude of 800km. The eccentricity is ε= A detailed description of the operational orbit can be found in the Mission Analysis chapter. The coronagraph science dedicated orbits are split into: a 12h long Apogee Pass, 1h for Breaking the Formation after the apogee pass, a 10h Perigee Pass, and 1h for Building the Formation again after the perigee pass. For the FF demonstration orbits the Perigee Pass part of the orbit is defined for the true anomaly ν=±137deg at which angle the altitude is 27022km. The duration of the perigee pass, t pp, is t pp =14080s (1)

77 Page 75 of 231 and the experimental part is t ff =849920s (2) or approximately about 4h and 20h respectively. During the Perigee Pass, in both cases, no forced motion relative manoeuvres are performed, but the satellites are still under closed loop control Coordinates frames We are using the following coordinate system for all the computations and the results Local Orbital Frame F o : Has its origin at the centre of mass of the passive spacecraft and the axes are defined as follow. This frame is also often referred to as the Local Vertical Local Horizontal (LVLH) frame. For the target spacecraft this frame will be referred to as F t and for the chaser spacecraft as F c. X o -axis: X o =Y o Z o which is in the direction of the velocity vector of the spacecraft. In the Rendezvous literature it is often referred to as the V-bar. Y o -axis: normal to the orbital plane and in the opposite direction and parallel to the orbital angular momentum vector. In the Rendezvous literature it is often referred to as the H- bar. Z o -axis: in the orbital plane from the spacecraft COM towards the Earth centre. In the Rendezvous literature it is often referred to as the R-bar. This rotating frame is selected in order to derive a linear system of differential equations whose representation and solution is tractable. The representation, in an inertial frame, becomes prohibitively complex Relative dynamics The relative dynamics for the classical Rendezvous and Docking (RVD) for quasi circular orbits is described by the Hill equations and solved by the very well known Clohessy-Wiltshire equations. They are nevertheless not applicable for any eccentric orbit. The relative dynamics is based on the Linear Time Varying (LTV) differential equations and the general closed form solution derived in RD[19]. The forces needed to SK at a certain location in the Local Vertical Local Horizontal (LVLH) frame are derived in equation 3 F x =m(kω 3 2x 0 -ω & z 0 -ω 2 x 0 ) (3) F y =mkω 3 2y 0 (4) F z =m(ω & x 0 -ω 2 z 0-2kω 3 2z 0 ) (5) where the index 0 means the station keeping location F i the force and m the spacecraft mass. ω is the true anomaly rate and k a constant for this time varying system. For the GNC computation a

78 Page 76 of 231 mass of 150kg has been used and an inertia of I=15kg m 2, which differs from the final mass but will not affect the accelerations given. It shall also be noted that the HEO on one hand gives a less externally perturbed orbit, but that the relative dynamics is an LTV system. This adds significant complexity for the design of the controllers as well as the guaranteeing stability margins. This is unproven ground for such a system, with very few tools available for the GNC designers, a point which should not be overlooked. 9.3 Manoeuvres The manoeuvres will be split up into the FF demonstration manoeuvres, followed by the ones needed for the coronagraph science part of the mission FF demonstration manoeuvres The Formation Flying manoeuvres are split up as follows. They are of generic nature, but it is ensured that they cover the most likely customer missions, namely Xeus and Darwin. Deployment and formation creation: This part is of both operational and demonstration nature, as it is needed by the mission itself. It is not known at which distance the spacecraft will come to rest, so as an average let us take half the max distance. (4km, y, z and 8km x) Formation station keeping on RF: Station keeping is performed at the prescribed manoeuvres. The far reach manoeuvre for RF will for fuel savings reasons be performed only along the x-axis. (25, 250m (x, y, z), 8km(x)) Formation station keeping on LM: Same profile as for RF but to a much higher precision. (25, 250m (x, y, z)) Formation keeping standby: This is just intended as a low fuel consuming waiting position in between manoeuvres on the RF or LM. (25m, x-axis) Formation resize: This will perform linear manoeuvres from 25m to 250m and back. It will be performed as a minimum time motion with full acceleration followed by deceleration at the mid point. This along the 3 axes. (25-250m, x, y, z) Formation in plane rotation: This performs a rotation about the axis normal to the plane of the prescribed angle forth and back at 2 different radii. (25-250m, 120deg in xz, xy, yz plane) Simultaneous rotation and resize: This will perform the previous 2 manoeuvres simultaneously (as above). Formation slew to retarget: This can be considered covered when rotations in all 3 planes have been performed. An additional distance of 1km is needed to cover slew for the last bullet (large angle). Formation station keeping for the EMMA configuration: This will cover the SK at a 1km distance as well as a manoeuvre to get there and back. (1km) Coronagraph Science manoeuvres This part will cover the coronagraph science manoeuvres, which consist of two separate ones as follows:

79 Page 77 of 231 Solar Observation: This observation phase is on the orbital arc, which is between ν=163deg and ν=197deg. The chaser spacecraft is at a relative distance of 120m along the line from the passive spacecraft origin to the sun. 2-Pulse Perigee Pass Manoeuvre: This is performed by a classical V manoeuvre with 2 pulses. This will bring the spacecraft from its final inertial pointed location to the new similar one on the other side GNC Requirements The GNC will have to handle all manoeuvres for both translation and rotation for all parts of the mission. The coronagraph science part of the mission is not the driver, the FF demonstration manoeuvres and the optical metrology are, where the GNC relevant criteria are listed below: Pos. Vel. Att. Rel. Att. Range Deployment [0.5;44] m 0.1 m/s 1 deg 1 deg 0.5 m -> 8 km SK RF 1 cm 1 mm/s 1 deg 1 deg 25, 250, 8000 m SK LM 32 µm 0.3 mm/s N/A 13.6 as 25,250 m Rotation In-plane 32 µm 0.3 mm/s N/A 13.6 as 25,250 m, 120 deg Resize 32 µm 0.3 mm/s N/A 13.6 as m Rotation Resize 32 µm 0.3 mm/s N/A 13.6 as m, 120 deg Slew 1 cm 0.3 mm/s N/A 13.6 as 25,250,1000 m, 120 deg Emma 1 mm 0.3 mm/s N/A 13.6 as 1000 m, 120 deg 9.4 V Computations Table 9-1: Criteria for the GNC system The calculations are split up into the separate manoeuvres, which will have the identifier of the dedicated section. Maximum accelerations are for the orbital driven contribution and the true anomaly where it happens. The accelerations for the linear manoeuvre part are not given as known from design Assumptions Perturbations and closed loop control is not taken into account, so we deal with a guidance problem. For rotations, centripetal accelerations are taken into account together with the orbital part and the same for all linear manoeuvres. The profiles are considered as minimum time manoeuvres leading to full acceleration followed by braking, with no constant speed part Station Keeping at 25m x-axis This provides the values for a station keeping for 1 hour at the given location. It shall be noted that the computation is performed at the ν=137deg. If it was performed at apogee it would be several times less, but already at 137deg the results are low. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e

80 Page 78 of Station Keeping at 25m y-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Station Keeping at 25m z-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Station Keeping at 250m x-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Station Keeping at 250m y-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Station Keeping at 250m z-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Station Keeping at 8km x-axis This is performed in the neighbourhood of the apogee in order to save fuel. It is about 10 times cheaper than at the beginning of the operational part of the orbit. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Station Keeping at 25m x-axis Standby This is to give an indication of the cost of a standby position for some time in closed loop. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e

81 Page 79 of Resize m along x-axis It shall be noted that the vast majority of the V comes from the manoeuvre linear motion and not from the orbital dynamics. Values are calculated for one way. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Resize m along y-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Resize m along z-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Resize m along x-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Resize 1-8km along x-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Rotation of 120deg in xz-plane at 25m The rotation takes place with a constant angular velocity, so here there is a cruise phase. The values are for one way. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation of 120deg in xy-plane at 25m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e+03

82 Page 80 of Rotation of 120deg in yz-plane at 25m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation of 120deg in xz-plane at 250m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation of 120deg in xy-plane at 250m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation of 120deg in yz-plane at 250m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation/Resize of 120deg in xz-plane at m Here the rotation and resizing are simultaneous, starting from the small radius. The values are for one way. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation/Resize of 120deg in xy-plane at m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e Rotation/Resize of 120deg in yz-plane at m V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e+03

83 Page 81 of Station Keeping Emma at 1km x-axis V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e Coronagraph Science observation As an example, the results of the case where the Sun vector is in the orbital plane and parallel to the semi minor axis are shown below. Over the mission these values will vary a bit as the orbit moves with respect to Sun, but not so much that the results are not representative. V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e The total V needed for the coronagraph science observation, using the above values, for the mission duration of 19 months is m/s Pulse Perigee manoeuvre The values given here are for ending and starting a coronagraph science cycle, so inertial pointing and for the case in Section V[m/s] Max. acc [m/s 2 ] ν[deg] Duration [s] e e e e+04 In Figure 9-2, the relative trajectory of the perigee pass manoeuvre for the case in Section is illustrated. The phase-plane plots are also shown to illustrate the relative velocities.

84 Page 82 of 231 Figure 9-2: Illustration of a typical perigee pass 2 pulse manoeuvre of trajectory in the x-z plane as well as the respective phase-plane plots Sequencing In order to provide a realistic view of the manoeuvre consumption, a first attempt to concatenate the manoeuvres is seen in the tables below. Each table is one orbit, recalling that the section number of the manoeuvre is the identifier tag.

85 Page 83 of 231 Orbit 1: Section Description Station Keeping at 25m x-axis Station Keeping at 25m x-axis Resize m along x-axis Station Keeping at 250m x-axis Station Keeping at 250m x-axis Rotation of 120deg in xz-plane at 250m Rotation of 120deg in xz-plane at 250m Rotation of 120deg in xz-plane at 250m Rotation of 120deg in xz-plane at 250m Station Keeping at 250m x-axis Station Keeping at 250m x-axis Resize m along x-axis Station Keeping at 25m x-axis Station Keeping at 25m x-axis Station Keeping at 25m x-axis Standby Pulse Perigee Table 9-2: Sequencing of orbit 1 Orbit 2: Section Description Station Keeping at 25m y-axis Station Keeping at 25m y-axis Resize m along y-axis Station Keeping at 250m y-axis Station Keeping at 250m y-axis Rotation of 120deg in xy-plane at 250m Rotation of 120deg in xy-plane at 250m Rotation of 120deg in xy-plane at 250m Rotation of 120deg in xy-plane at 250m Station Keeping at 250m y-axis Station Keeping at 250m y-axis Resize m along y-axis Station Keeping at 25m y-axis Station Keeping at 25m y-axis Station Keeping at 25m x-axis Standby Pulse Perigee Table 9-3: Sequencing of orbit 2

86 Page 84 of 231 Orbit 3: Section Description Station Keeping at 25m z-axis Station Keeping at 25m z-axis Resize m along z-axis Station Keeping at 250m z-axis Station Keeping at 250m z-axis Rotation of 120deg in yz-plane at 250m Rotation of 120deg in yz-plane at 250m Rotation of 120deg in yz-plane at 250m Rotation of 120deg in yz-plane at 250m Station Keeping at 250m z-axis Station Keeping at 250m z-axis Resize m along z-axis Station Keeping at 25m z-axis Station Keeping at 25m z-axis Station Keeping at 25m x-axis Standby Pulse Perigee Table 9-4: Sequencing of orbit 3 Orbit 4: Section Description Resize m along x-axis Station Keeping Emma at 1km x-axis Station Keeping Emma at 1km x-axis Resize 1-8km along x-axis Station Keeping at 8km x-axis Station Keeping at 8km x-axis Station Keeping at 8km x-axis Resize 1-8km along x-axis Station Keeping Emma at 1km x-axis Station Keeping Emma at 1km x-axis Resize m along x-axis Station Keeping at 25m x-axis Standby Pulse Perigee Table 9-5: Sequencing of orbit 4

87 Page 85 of 231 Orbit 5: Section Description Rotation of 120deg in xz-plane at 25m Rotation of 120deg in xz-plane at 25m Rotation of 120deg in xy-plane at 25m Rotation of 120deg in xy-plane at 25m Rotation of 120deg in yz-plane at 25m Rotation of 120deg in yz-plane at 25m Station Keeping at 25m x-axis Standby Pulse Perigee Table 9-6: Sequencing of orbit 5 Orbit 6: Section Description Rotation/Resize of 120deg in xz-plane at m Rotation/Resize of 120deg in xz-plane at m Rotation/Resize of 120deg in xy-plane at m Rotation/Resize of 120deg in xy-plane at m Rotation/Resize of 120deg in yz-plane at m Rotation/Resize of 120deg in yz-plane at m Station Keeping at 25m x-axis Standby Pulse Perigee Table 9-7: Sequencing of orbit Summary of V computations For the formation flying manoeuvres, a very minimum of 6 orbits to perform them once is needed. It is deemed necessary to perform this several times, and as a strict minimum 2 times, for which the following values are based. To the result is added a 15% margin to cover the perturbations and closed loop control. FF demonstration Manoeuvres [m/s] The coronagraph science part is computed for a period of one year, with a 12h apogee pass, taking the same margins as for the previous table Coronagraph Science Observations [m/s] ~ Number of Satellites In order to perform a proper and useful FF demonstration, the number of satellites is very important. 3 or more satellites are basically needed to perform the manoeuvres needed by, for

88 Page 86 of 231 example, the Darwin mission and in order to span a plane in space. This cannot be performed with only 2 satellites, which falls back in the RVD case. This is illustrated in Figure 9-3. Figure 9-3: Here is illustrated a 3 satellite-formation and what the manoeuvres relate to Three satellites option Figure 9-3 shows 3 satellites and what it means to resize, rotate and both simultaneously. It is important to realize that the 3 span a plane in space. This means the degrees of freedom are all locked. This gives cross couplings in the system, which is a very important point to demonstrate at formation level. The variables a and b are the reference locations and d 1 and d 2 the control deviation with respect to those. We can write the equation of a+d 1 =k(b+d 2 ), where k is some scale factor different from zero. A core point in the GNC demonstration is that when a+d 1 is varied, the other satellite referring to k(b+d 2 ), shall automatically be controlled to fulfil the relation. The variable c is used as a triangularization measurement to obtain an independent measure of satellite relative motion to both the central one and between them Two satellites solution (Baseline) In the case of having only 2 satellites, which are the encircled ones in Figure 9-3, it is obvious that the coupled relationship between satellites simply breaks down, and cannot be demonstrated. Further to that, there is also no longer the span of a plane in space. This means that the configuration falls back to the classical 2 satellite Rendezvous case. This also leaves one axis rotational degree free, with no coupling to position as it would have in the 3 satellites case. Such scenarios have been in operation in space for many years, though with less accuracy.

89 Page 87 of 231 It leaves the situation that few of the demonstration manoeuvres and systems are novel, only simultaneous resize and rotation, whereas the rest have been performed Relation to other European missions Two other European missions are performing similar GNC relative manoeuvres, as follows: The Prisma mission is a two spacecraft LEO demonstrator, basically for RVD. It will be flight testing the same RF metrology system as in. This can either be regarded as duplication or independent verification. Many manoeuvres are similar in the two missions. The other project is the Automated Transfer Vehicle (ATV), that is the European supply vehicle for RVD with the International Space Station. The ATV will basically perform most of the manoeuvres discussed here for, but in the centimetre accuracy range. Careful consideration should be paid to the complementary manoeuvres of the mission under study, with respect to the Prisma and ATV missions, in particular when considering the two satellite formation. This in order to ensure the best return of the demonstration mission Demonstration coverage of customer missions With respect to the demonstration coverage of the two most likely FF missions Darwin and Xeus, it is judged that the subject demonstration mission would have approximately the following coverage: Darwin Xeus # spacecraft Manoeuvres 20 % 100 % 100 % Precision 25 % 80 % 60 % Table 9-8: The demonstration coverage of customer missions 9.6 Avionics Equipment for FF The equipment needed by the GNC is detailed in the following sections and the equipment proper in their respective chapters Propulsion and pulse width modulation The propulsion selected is a cold gas system in order to fulfil the overall needs of the mission. The thrusters are on/off type which will be Pulse Width Modulated (PWM) in order to obtain an effectively smaller equivalent thrust. The computations are performed such that it shall be possible to provide an acceleration and breaking pulse with a duration of a sampling time T s and stay within the requirements of 32µm. The velocity must be smaller than 0.3mm/s. Based upon typical GNC closed loop bandwidths from the ICC study and that there are no drivers for fast manoeuvres, we select T s =10s. We share the values evenly between translation and the translation contribution from the rotation, considering a lever arm of 0.5m. The needed amplitude of the force can be found from

90 Page 88 of 231 ms 2 F= T 2 s (6) where s 2 is the width of the requirement. A similar relationship exists for the rotation. With the values above this leads us to the require forces F reg 24µN (7) The driver is the position requirement and neither the velocity nor the relative attitude requirement. To achieve the maximum thrust required for the manoeuvres thrusters of level 10mN and 40mN is foreseen. They have a Minimum Impulse Bit (MIB) of MIB 10 = 70µNs (8) and MIB 40 = 260µNs (9) and we therefore can safely use a minimum on time of T onmin =10ms (10) This means that with PWM we are able to achieve a minimum thrust of F min 10µN (11) which is smaller than that required in equation 7. The thrusters have a settling time of less than 10ms and a 5% noise of nominal thrust at steady state and 15% during rise, both at 3σ. In order to obtain clean force and torque capability, 12 sets of the two thrusters are needed. They are mounted with two on each surface of the spacecraft and in the 3 planes of the spacecraft body frame, with the force vectors parallel to the body frame axes. It shall be observed that a smaller number can be used with some sort of canted nozzles, through introducing couplings into the control system. Such analysis and design is nevertheless beyond the scope of this study activity Star tracker In order to fulfil the demonstration with the above requirements, a star tracker is needed which can provide an output on 2 axes with an accuracy better than 0.55arcsec. This could probably be relaxed to about 1-1.5arcsec, as the GNC provides better performance than the requirement. All values 3σ. A possible candidate for such equipment could be the DTU star camera, which is expected, in the near future, to be able to deliver arcsec accuracy. It has a mass of 1kg and a power consumption of 3.8W.

91 Page 89 of Gyros Gyros are needed for the deployment phase and detumbling, as well as during the manoeuvres for the attitude control. In order to be in line with the requirements and the contributions from position and attitude cross couplings, the performance need be better than deg/s at 3σ. This can, for example, be provided by the SiRRS01-05 MEMS from BAE, which has a drift of 3 deg/h. It has a mass of 0.6kg and a power consumption of 3.5W Coarse Sun sensor These sensors are needed for the safe mode and also deployment. They shall be used such that there is 4π steradian coverage. An accuracy of a few degrees is sufficient. A candidate for this is the new TNO Micro Digital Sun sensor, all in a small box sensor expected to be available in 1-2 years. It has a mass of about 80g and no power from the bus as self supplied wireless Bluetooth communication RF metrology For details and performance of this system under test, please see Chapter 6. This is an essential system for any part of the mission Optical metrology The full package of the laser metrology system is needed (HPOM), except for the scanning interferometer. The reason for the latter being left out is that nanometre accuracy cannot be achieved. For details and performance of this system under test please see Chapter 7.

92 Page 90 of 231 This Page Intentionally Blank

93 Page 91 of CONFIGURATION 10.1 Requirements and Design Drivers The configuration has to comply with the following requirements: Shall fit inside the fairing of the VEGA launcher and the Eurockot launcher Shall accommodate all equipment and instruments Shall provide unobstructed fields of view for the coronagraph instrument, the optical metrology and the antennas Shall provide access to install and service components during ground operations. The design driver was to keep the configuration straightforward and as small as possible, such that the structural mass would be minimal Assumptions and Trade-Offs At the beginning of the study, a trade-off was made between a box shaped structure and a cylindrical shaped structure for the spacecraft. The box structure was chosen because it is easier to accommodate and attach the equipment and instruments to the flat surfaces of the box structure. A second reason was that, at the beginning of the study, the baseline launcher was Eurockot with its Breeze Upper Stage. At its interface, Breeze has a standard pattern of attachment holes. The idea was to attach the four load carrying corner stringers of the box structure directly to these discrete interface points. Later in the study, the LISA Pathfinder Propulsion Module (LPF PM) was introduced, to be mounted between the launcher and the spacecraft. A drawback of the chosen box structure, in this new stacked configuration, is the more complicated and heavier attachment to the circumferential interface of the LPF PM Two Satellites Solution (Baseline) In Figure 10-1, the two spacecraft solution is shown stacked inside the VEGA fairing on the Launch Vehicle Adapter (ACU 937). The completely stacked configuration from top to bottom consists of: 1. Coronagraph satellite 2. Occulter satellite 3. Adapter between the LPF PM and the Occulter satellite 4. LPF PM (short version) with a length of about 2061 mm, bottom diameter 900 mm and top diameter 800 mm (LPF PM is still under development). The total height of the stack is 4250 mm and from the image, it is clear that the stack fits well inside the 2380 mm diameter usable volume envelope of the VEGA fairing. The complete stack fits in the fairing of Eurockot launcher as well.

94 Page 92 of 231 Coronagraph Satellite Occulter Satellite Adapter LISA Pathfinder Propulsion Module Launch Vehicle Adapter (ACU 937) Figure 10-1: Stacked launch configuration inside VEGA fairing

95 Page 93 of Coronagraph satellite The global dimensions of the Coronagraph satellite are 756 x 900 x 850 mm 3 (see Figure 10-2) and the dimension of each solar panel is 1000 x 750 mm 2. Figure 10-2: Dimensions of the Coronagraph spacecraft The main instrument carried on this satellite is the coronagraph which consists of a telescope (900 x 300 x 300 mm 3 ) and an electronic box (200 x 200 x 200 mm 3 ). For the Optical Metrology Subsystem there are two elements on the Coronagraph satellite: a Fine Lateral Sensor transmitter and a Corner-Cube Retro-reflector. The locations of these elements are shown in Figure Furthermore, there are two apertures shown in the front panel of the satellite; one for the Coronagraph telescope and one for the Corner-Cube Retro-reflector.

96 Page 94 of 231 Coronagraph Telescope Electronic Box Coronagraph Corner-Cube Retro-reflector Fine Lateral Sensor transmitter Figure 10-3: Coronagraph instrument and elements of the Optical Metrology Subsystem on the Coronagraph satellite In Figure 10-4, the Coronagraph instrument and all elements of the Optical Metrology Subsystem are hidden, such that the elements of the Radio Frequency Metrology Subsystem and the elements of the Communication Subsystem can better be seen. On the front panel and on the back panel as well (not shown) there are three S-band LGA antennas for FF and one S-band LGA antenna for TT & C. On the bottom panel (and on the top panel as well) there are one GPS antenna and one S-band LGA antenna for TT & C. All electronic boxes for the RF FF Subsystem and Communication Subsystem are accommodated on the same side panel.

97 Page 95 of 231 TM/TC Unit (S-band Transceivers) S-band LGA (TT & C) S-band LGA (FF) GPS Receiver RF FF Electronic Box S-band LGA (FF) GPS Antenna S-band LGA (Formation Flying) S-band LGA (TT & C) Figure 10-4: Radio Frequency Metrology elements and Communication Subsystem elements In Figure 10-5, the two sun sensors (on the front and the back panels), four star-trackers (on the top and the bottom panels) and four reaction wheels (on one side panel) are shown. The blue sphere in the middle is the tank with the propellant for the twenty-four small thrusters. The green boxes are electronic boxes of the Data Handling Subsystem and the red boxes are the battery and the PCDU of the Power Subsystem.

98 Page 96 of 231 Figure 10-5: Other elements inside the Coronagraph Satellite

99 Page 97 of Occulter satellite The global dimensions of the Occulter satellite are 756 x 1005 x 850 mm 3 (see Figure 10-6) and the diameter of the Occulter disc, covered with solar cells, is 1200 mm. Figure 10-6: Dimensions of the Occulter satellite Besides the solar panel and the bigger tank, the elements of the Optical Metrology Subsystem are different compared to the Coronagraph satellite (see Figure 10-7). All other subsystems of the Occulter satellite are similar with the subsystems of the Coronagraph satellite.

100 Page 98 of 231 Fringe Tracking Sensor Coarse Lateral Sensor Absolute Longitudinal Sensor Absolute Longitudinal Sensor Head Fine Lateral Sensor Receiver Figure 10-7: Elements of the Optical Metrology Subsystem

101 Page 99 of PROPULSION 11.1 Requirements and Design Drivers The propulsion subsystem to be used on is mainly dedicated to providing orbit correction and fine attitude control during the formation flying manoeuvres. A list of requirements drives the selection of the propulsion system to be used. They can be identified as: ORBIT SELECTION. HEO with perigee altitude of 800km, eccentricity of 0.829, inclination of 60 VEGA/VERTA Launcher (63 EUROCKOT/DNEPR option) Consequent to the orbit selection and considering FF manoeuvres requirements, reduced compare to DARWIN requirements, the THRUST RANGE (required by GNC) is in the order of: o Tens of mn (in the order of 40mN) for FF manoeuvring o Tens of µn for fine attitude control and FF manoeuvring. 6DoF on each satellite (required by GNC). The propulsion system has to include a minimum set of 12 thrusters on board of each satellite Schedule. Launch date 2009/2010 Power budget limited Mass of each satellite, in the 200kg range (micro-satellite) Size of each satellite, in the 800x900x900mm range (micro-satellite) Assumptions and Trade-Offs A trade off among different technologies has been performed. EP Systems and Cold Gas systems have been considered. A table at the end of the chapter summarizes the TRL for each Thruster Electric propulsion trade off The following thrusters have been considered: RIT 10, developed by EADS ST, Flight Proven in ARTEMIS, candidate for DARWIN T5, developed by QinetiQ, Flight Proven in ARTEMIS, FM for GOCE, candidate for DARWIN ALTA 150 µn-feep, developed by ALTA, EM, selected for MICROSCOPE, candidate for LISA PF, LISA, GAIA, DARWIN In FEEP, developed by ARC, EM, candidate for LISA PF, LISA, GAIA, DARWIN RIT 4, developed by GIESSEN University, Prototype, candidate for DARWIN Mini HET, developed by ALTA, EM, candidate for DARWIN QINETIQ mini Ion Engine, candidate for DARWIN RMT, developed by ALCATEL ALENIA SPACE (LABEN/PROEL), EM, candidate for DARWIN For each Thruster the performances and the system mass budget and volume are given (values given for 12 thrusters).

102 Page 100 of RIT 10 After over 25 years of extensive research and development, the first Radio-frequency Ion Thruster Assembly (RITA) was successfully demonstrated in space onboard the ESA's European Retrievable Carrier (EURECA), launched by the Space Shuttle Atlantis in At that time, the RIT-10 system provided a nominal specific impulse of 3058 seconds. More recently, RITA-10 was used to retrieve the ARTEMIS satellite from a total loss to a full recovery, after thrusting for 6430 hours. RITA-10 is space qualified and has demonstrated thousands of hours of flawless operation in space and tens of thousands of hours on ground. Figure 11-1: RIT10 Thruster Unit Figure 11-2: T5 & RIT10 on ARTEMIS Each Radiofrequency Ion Thruster Assembly (RITA) consists of the following units: 1 Flow Control Unit (FCU) 1 RIT 10 Thruster and neutralizer 1 Radio Frequency Generator (RFG) 1 Power Supply and Control Unit (PSCU)

103 Page 101 of 231 Figure 11-3: RITA Operation principle The performance of the RIT 10 Thruster is listed in the following table. Performance Nominal Demonstrated Thrust 15mN 1-41mN (0.33mN minimum for the RIT10 EVO) Thrust resolution 80µN 1µN (for the RIT10 EVO) Specific Impulse 3300s s Total Impulse 0.81E6 N.s (Artemis Mission) 1.07E6 N.s On Time 15,000 Hrs (Artemis Mission) 20,083 Hrs Cycles 5,000 (Artemis Mission) 6,741 Thrust Vector Stability < ± 1 Beam Divergence 11.5 Specific Power 27.5W/mN Table 11-1: RIT 10 performance The mass budget and the volume data are available in the following tables.

104 Page 102 of 231 Table 11-2: RITA system mass budget Table 11-3: RITA system volume Reasons for not being selected as baseline: The mass budget calculation, due to the requirement of having 12 thrusters onboard, leads to a heavy propulsion system. An exercise could be done in order to reduce the numbers of PSCU and RFG in the case that not all the thrusters are demanded to operate at the same time. Moreover, since a higher thrust level (in the order of 40mN) is demanded, more power will be required T-5 The T-5 Thruster was developed from the UK-10 and UK-25 thrusters using well-established and proven scaling laws. A larger T-6 thruster is currently under development for future Telecommunication and interplanetary missions. All of these thrusters have exhibited good performance, resulting principally from refined ion optics and efficient discharge chamber designs. The mechanical design of the T-5 thruster is the product of three decades of research and development. During this time, the basic plasma and ion beam features have remained virtually unchanged. As a consequence, the majority of the test data from this long period became applicable to scientific missions such as GOCE. An engineering model (EM) T-5 has been successfully tested, for more than 5000 hours, in support of the GOCE program. In contrast to Telecommunication applications (for example ARTEMIS/ALPHABUS) the GOCE mission will demonstrate the capability of achieving wide throttle range and stringent thrust control requirements with Ion Engine Technologies. Each Electron Ion-Bombardment Thruster Assembly (EITA) consists of the following units: 1 Power Supply and Control Unit (PSCU) 1 Proportional Xenon Feed Assembly (PXFA) 1 T-5 thruster with neutralizer

105 Page 103 of 231 Thruster Propellant Propellant Inlet S/C Power EIT HV Power + Return IPCU Figure 11-4: EITA system The performance of the T-5 thruster is listed in the following table. Table 11-4: T5 performance The mass budget and the volume data are available in the following tables.

106 Page 104 of 231 Table 11-5: EITA system mass budget Table 11-6: EITA system volume Reasons for not being selected as baseline: The mass budget calculation, due to the requirement of having 12 thrusters on board, leads to a heavy propulsion system. An exercise was done in order to reduce the mass budget considering that the biggest contribution comes from the PSCU. It was considered: to use 12 PSCU to use 6 PSCU, in the case half of the thrusters need to operate at any one time to use a distributed power supply with a single high voltage source. The numbers in Table 11-5 refer to the latter one. Since a higher thrust level (in the order of 40mN) is demanded, more power will be required Alta 150 µn-feep Field Emission Electric Propulsion (FEEP) is currently the object of great interest due to its unique features:

107 Page 105 of 231 sub-µn thrust range, ( µN) near instantaneous switch on/switch off capability high resolution (0.1µN) high specific impulse (from 5000 up to 7500s), that is, very low consumption of propellant. Since 2001, Alta has been responsible for the MICROSCOPE micro-propulsion subsystem definition and development under ESA contract. Since 2002, Alta is also working on similar subsystem elements to be flown as technology demonstrators in preparation of the ESA LISA Pathfinder mission. In 2003 the pre-development activities were extended, under ESA contract, to cover also the GAIA mission requirements. The FEEP-150 is also a candidate for DARWIN. The FEEP-150 propulsion subsystem is presently being qualified by Alta for the CNES Myriade platform (MICROSCOPE) under ESA contract. Subcontractors are: Galileo Avionica (with responsibility for the high voltage power supply and emitter micromachining), Alcatel Alenia Space (with responsibility for the neutralizer), Contraves Space (with responsibility for the sealing and opening mechanism) and EADS Astrium (with responsibility for surface tension tank verification). The FEEP-150 sub-system is made up of self-contained thruster clusters (EPSAs). Each EPSA includes: 3 Thruster Assemblies (TA), including thruster unit and propellant tank 1 Power Processing and Control Unit (PPCU) 2 Neutralizers (NA), one active and one cold redundant. Figure 11-5: FEEP-150 TA

108 Page 106 of 231 Figure 11-6: MICROSCOPE FEEP-150 EPSA The performance of the FEEP-150 Thruster is listed in the following table. Table 11-7: FEEP-150 performance The mass budget and the volume data are available in the following tables. Table 11-8: EPSA mass budget Table 11-9: EPSA volume Reasons for not being selected as baseline: Thrust level required. The power consumption.

109 Page 107 of ARC Indium FEEP The thruster is based on the space-proven ARC Indium Liquid-Metal-Ion-Source technology. This technology is the core of active spacecraft potential control devices of mass spectrometer instruments flying on a number of satellites including CLUSTER-II, ROSETTA and DOUBLESTAR (accumulated over 6600h in space). Each emitter can produce up to 100µA, which translates into thrust of up to 10µN. Depending on the acceleration voltage, the specific impulse ranges roughly between 4000 and 8000s. The emitters can be clustered to obtain higher maximum thrust levels. An endurance test of 5000h has already been performed. The In-FEEP is candidate for LISA PATHFINDER, GAIA and DARWIN. Each module proposed for LISA PATHFINDER includes: 16 In-FEEP emitters, each one with an Indium reservoir of 14g (max thrust of 200µN) Integrated pre-resistor for single power supply operation 1 Power Supply Neutralizer Figure 11-7: ARC In-FEEP The performance of the in-feep is listed in the following table. Table 11-10: In-FEEP performance The mass budget and the volume data are available in the following tables.

110 Page 108 of 231 Table 11-11: In-FEEP system mass budget Reasons for not being selected as baseline: Tens of mn required. The power consumption Alta Mini HET The miniaturized Hall Effect Thruster (mini-het) is a small, low-power HET designed to perform orbit control tasks on micro-satellites and attitude control tasks on mini-satellites. The system is particularly interesting for future ESA Science Missions such as DARWIN and Earth Observation Mission (Optical Mission, LIDAR Mission, InSAR Formation Flying Mission and Earth Gravity Mission). The thruster unit is fully based on European know-how and technology, and the same applies to all the key sub-system components. Operated on % pure xenon, the Thruster provides a specific impulse ranging between 900s and 1700s, with power ranging from 50W to 250W and thrust level between 0.5 and 10mN. The mini-het subsystem is designed to have minimum system level impact when integrated on small satellites. The subsystem is built around one or more thruster units, and can be supplied with different options for what concerns cathode, power supply, propellant control, tank and (if needed) a pointing mechanism and a diagnostic package. The mini-hall Effect Thruster System includes: HETs Neutralizers Power Supply & Control System (PSCU) able to drives up to 4 thrusters Xenon Flow Control System (XFCU). Figure 11-8: ALTA mini HET The performance of the mini- HET is listed in the following table.

111 Page 109 of 231 Table 11-12: ALTA mini-het performance The mass budget and the volume data are available in the following tables. Table 11-13: ALTA mini-het system mass budget Table 11-14: ALTA mini-het system volume Reasons for not being selected as baseline: Minimum thrust level required for formation flying manoeuvres (order of 10uN). The power consumption Mini Radiofrequency Ion Thruster: RIT4 As a response to several scientific missions asking for a thrust range in the µn level, a study on the process of scaling down the size of the RIT thruster has been initiated in the Giessen University (Germany). Two Laboratory Prototypes, the µn-rit4 and the µn-rit2, have already been manufactured and tested. The advantages of the mini RITs, besides their low thrust range, are their controllability and the mass savings in long-term missions due to their reduced size and

112 Page 110 of 231 the higher specific impulse. The usage of inert gas, Xenon, avoids possible S/C contamination. Furthermore, within the class of gas-discharge electric thrusters, the radiofrequency type seems to be the most suitable to scale down, because it works without any discharge electrodes, and the magnetic pole is inside the ionizer. The absence of discharge electrodes also allows for a long lifetime. Another advantage of the mini RIT is its regulation by varying gas flow, radiofrequency power and extraction voltages independently. The mini RITA consists of the following units: 1 RIT 4 Thruster and neutralizer 1 Flow Control Unit (FCU) 1 Radio Frequency Generator (RFG) 1 Power Supply and Control Unit (PSCU) Figure 11-9: RIT4 The performance of the mini- RIT is listed in the following table. Table 11-15: RIT4 performance The mass budget and the volume data are available in the following tables.

113 Page 111 of 231 Table 11-16: RIT4 system mass budget Table 11-17: RIT4 system volume Reasons for not being selected as baseline: The maximum thrust level required (in the order of 40mN), the power consumption and the TRL RMT The RMT (Radiofrequency with Magnetic field ion Thruster) is a fine throttlable Ion Thruster in the milli-newton range (0-12mN), with a resolution of less than 0.1mN, exploiting high efficiency (overall power consumption less than 500W at max thrust) and high specific impulse ( s). The RMT has been developed in the frame of an R&D contract awarded by the Italian Space Agency for application of Drag Free Station Keeping tasks for Satellite platforms in the kg class or for Attitude Control tasks of Scientific Satellites. The Propulsion system is composed of the following units: Thruster Module (TM), Gas feed line & Flow Control Unit (GFCU) including the high pressure regulator control valve and low pressure thermo throttle flow chain Radiofrequency Generator & Matching network (RFGM) Power Supply & Control Unit (PSCU)

114 Page 112 of 231 Figure 11-10: RMT The performance of the RMT is listed in the following table. Table 11-18: RMT performance The mass budget and the volume data are available in the following tables. Table 11-19: RMT system mass budget

115 Page 113 of 231 Table 11-20: RMT system volume Reasons for not being selected as baseline: High thrust level required (in the order of 40mN). The power consumption Cold Gas thrusters trade-off The following cold gas thrusters have been considered: MAROTTA SV14 Cold Gas, FM for CRYOSAT Bradford Cold Gas, FM for GOCE Marotta Micro Cold Gas, EM candidate for GAIA Bradford Micro Cold Gas, EM candidate for GAIA Alenia Micro Cold Gas, BB/EM, candidate for GAIA Uppsala Micro Cold Gas, BB, candidate for GAIA. For each Thruster the performances and the system mass budget/volume are given Marotta SV14 Cold Gas The Marotta SV14 Cold Gas has been selected as baseline. For further information on the thrusters see Chapter Bradford Cold Gas The Bradford Cold Gas Thruster is the baseline for the GOCE Mission. Figure 11-11: Bradford Cold Gas The performance of the Bradford Cold Gas thruster is listed in the following table.

116 Page 114 of 231 Thrust range Specific Impulse 100µ to 50mN depending on design and input pressure >70s Table 11-21: Bradford Cold Gas performance Reasons for not being selected as baseline: Demanded thrust level. The Bradford Cold Gas Thruster has been qualified for the thrust range 400µN-600µN Marotta Micro Cold Gas The Marotta Micro Cold Gas Thruster is acandidate for the GAIA Mission. The EM of the Thruster is in preparation. Characterisations Test Campaigns have been conducted at Marotta UK and ESTEC-EPL. Environmental (Vibration, Shock, Thermal Cycling) and Life Testing (10 million Open/Close cycles) have already been performed. Figure 11-12: Marotta Micro Cold Gas The performance of the Marotta Micro Cold Gas Thruster is presented in the following figure.

117 Page 115 of 231 Figure 11-13: Marotta Micro Cold Gas performance Reasons for not being selected as baseline: Max Thrust required (in the order of 40mN). TRL Bradford Micro Cold Gas The Bradford Micro Cold Gas thruster is a candidate for the GAIA Mission. The solenoid valve was previously qualified for GOCE. Characterisations Test Campaigns have already been conducted at Bradford Engineering and ESTEC-EPL. Figure 11-14: Bradford Micro Cold Gas The performance of the Bradford Micro Cold Gas thruster is presented in the following figure.

118 Page 116 of 231 Figure 11-15: Bradford Micro Cold Gas performance Reasons for not being selected as baseline: Max Thrust required (in the order of 40mN). TRL ALENIA Micro Cold Gas The ALENIA Micro Cold Gas Thruster is currently a candidate for the GAIA Mission. To date several different BB/EMs have been manufactured (standard valve with nozzle, with heater for Xe use, standard pressure regulator). The mass flow sensor, used for Closed Loop feedback required further development. The thruster has been extensively tested (environmental testing included). The life test is on-going (1 million cycles completed to date). Figure 11-16: Alenia Micro Cold Gas Reasons for not being selected as baseline: Max Thrust required (in the order of 40mN). TRL Uppsala Micro Cold Gas The manufacturing of the BB hardware has been completed as part of a study in To date no validation or performance test results are available.

119 Page 117 of 231 Figure 11-17: Uppsala Micro Cold Gas Reasons for not being selected as baseline: Max Thrust required (in the order of 40mN). TRL TRL The TRL for the different thrusters (Electric and Cold Gas) are presented in Table The propulsion system chosen for the mission is indicated in red. It should be noted that for the T5 and the RIT10 a delta qualification is needed for, since the thrusters will be used in a pulsed mode and therefore the TRL is 7 in 2009 and 8 today. (*) different thrust level available according to size and inlet pressure Table 11-22: TRL

120 Page 118 of Two Satellites Solution (Baseline) The MAROTTA SV14 Cold Gas thruster has been selected as baseline for the propulsion subsystem to provide orbit correction and fine attitude control during the formation flying manoeuvres on the basis of: The demanded thrust levels (tens of mn, tens of µn) The requirement of having 12 thrusters on board (the envelope available) The maturity level, schedule The power availability. Two Thruster units are used to satisfy the thrust level required: Marotta SV14-001: Thrust=40mN, Isp=70s Marotta SV14-002: Thrust=10mN, Isp=70s. The thrusters are operated in on/off mode with a sampling time of 10s and a minimum ON time of 10ms. A total number of 24 cold gas thrusters are used on board of each satellite. The system chosen is the one flight qualified for the CRYOSAT mission (high TRL already reached; cycles life). The power required to operate each thruster is: 3.5W to open the valve and hold it for the first 75ms 0.7W for pulse durations of more than 75ms Coronagraph satellite The Coronagraph satellite has onboard: 12 x 40mN Cryosat Marotta Cold Gas thrusters 12 x 10mN Cryosat Marotta Cold Gas thrusters A pressurized tank (Ardé 30.8l), Ø420.12mm maximum outer shell envelope A feed system based on the Cryosat configuration. Figure 11-18: Marotta Cold Gas

121 Page 119 of 231 Figure 11-19: CRYOSAT feed system The maximum amount of propellant considered is 8kg (including a margin of 2%). The total mass allocated for the propulsion subsystem is 27.4kg. Table 11-23: Coronagraph mass budget Occulter satellite The Occulter satellite has on board: 12 x 40mN Cryosat Marotta Cold Gas thruster 12 x 10mN Cryosat Marotta Cold Gas thruster A pressurized tank (PSI 80l), Ø424mm X 752mm A feed system based on the Cryosat configuration. The maximum amount of propellant considered is 25kg (including a margin of 2%). The total mass allocated for the propulsion subsystem is 47kg.

122 Page 120 of Options Table 11-24: Occulter mass budget FEEP subsystem A FEEP subsystem (3 thruster units, 1 PPCU, 2(1) Neutralizer) is proposed to be embarked on each satellite as a payload (technology demonstrator). This allows the possibility to test a propulsion module having the resolution required by future ESA FF missions such as DARWIN Three Satellites Option For the three satellites option, the propulsion subsystem used for the third satellite (Free Flyer 2) is the same as that of the Occulter. A FEEP subsystem (3 thruster units, 1 PPCU, 2(1) Neutralizer) is proposed to be embarked on each satellite as a payload (technology demonstrator).

123 Page 121 of POWER 12.1 Requirements and Design Drivers The overall function of the power subsystem is to generate the required energy on board during the full mission duration. In addition, the power subsystem shall distribute the electrical energy on board to the different equipments through protected lines. Switching capabilities shall be implemented inside the power subsystem, enabling the reconfiguration of the bus in case of any failures. The power subsystem shall be tolerant of any single point failure. The main drivers of the power subsystem design are: Readiness of the technologies in line with the mission timeline The 19 months mission duration Cost and limitation of the required qualification/development processes To use European components as far as possible Mission overall requirements For power subsystem sizing purposes, the maximum eclipse duration for the selected HEO is assumed to be 1.3 hours. Additional details are in the Mission Analysis chapter. Both coronagraph science and FF demonstration orbits are detailed in the two following sections. As well as these two operational sequences, a three hours safe mode is also taken into account for the design of the space segment. In case of a failure of the attitude control, no power generation can be guaranteed by the solar cells and the battery will then have to supply the full power required on the bus Coronagraph Science Orbit A coronagraph science orbit is dedicated to Sun observation with the coronagraph instrument. While observing, the coronagraph has to be oriented towards the sun, with the occulter disc 120m in front, also pointing towards the Sun. In this configuration, the shadow of the 1.2m diameter shield mounted on the Occulter satellite results in a full shadowed area limited to the coronagraph instrument s optical part. Moreover, as illustrated in Figure 12-2 and Figure 12-3, the Coronagraph satellite is then in a partially shadowed zone where only a limited amount of solar flux can reach the satellite. The power generated by a cell will be directly linked to the distance from the satellite axis (where the coronagraph instrument is mounted).

124 Page 122 of 231 Figure 12-1: Coronagraph Science Orbit Phases Figure 12-1 shows a typical time breakdown of a fully dedicated coronagraph science orbit: At the apogee, the coronagraph will observe the Sun for up to 6 hours Around this period, the two satellites will remain with the same attitude and only communication will be performed. A maximum of 6 hours is allocated for that purpose For the remaining time of the orbit, the satellites are also expected to be sun pointed. Contrary to the coronagraph science mode, the Coronagraph satellite is not kept in the (partial) shadow of the Occulter, to allow a full use of the photovoltaic cells. Since the positioning of the two satellites for the Sun-corona observation is a complex task, where a Sun pointing attitude is difficult to maintain, two transient modes of one hour each with a maximum average Sun angle of 45 degrees are also assumed. Figure 12-2: Positioning of the satellites during coronagraph science mode

125 Page 123 of 231 Figure 12-3: Partial shadowing of the Coronagraph satellite (facial view) Along one coronagraph science orbit, the units will be activated or not depending on the operational modes. The states for the different units are illustrated in Table For instance, the RF metrology remains permanently ON while the optical metrology is only switched ON for the fine positioning during the 12 hours coronagraph science mode. Table 12-1: Coronagraph science mode: Equipment states during mission modes

126 Page 124 of FF demonstration Orbit When an orbit is dedicated to FF instead of coronagraph science, the sequence is entirely modified. Around the apogee, up to 20 hours is reserved to FF manoeuvres. Since the pointing during the manoeuvres is not determined in this study, the attitudes of both satellites with respect to the Sun are still undefined. Nevertheless, an attitude envelope will need to be defined in the Phase-A study in order to avoid unjustified over-design of the power subsystem. As for the Sun observation orbits, two periods of maximum one hour each are also allocated for the transition between the perigee passage and the FF demonstration segment of the orbit. As a sizing case, during these two periods, the average Sun depointing of the two satellites is assumed to stay under 45 degrees. During the remaining time of each orbit, excluding the potential eclipse period, both the satellites will be kept Sun-pointed in order to maximize power generation. Figure 12-4: FF demonstration Orbit For a better understanding of the FF demonstration orbits, Table 12-2 summarizes the operational states of the units of both satellites, in the identified segments.

127 Page 125 of 231 Table 12-2: FF: Equipment states during mission modes Power requirement on the bus The two satellites have the same subsystems (see the list in Table 12-3), except for the coronagraph payload and the optical metrology modules. In a configuration with three satellites, the third one would be identical to the Occulter satellite for cost and development reductions. The mission modes should also be unchanged compared to a two satellites configuration. The only identified change affecting the power bus is the implementation of the satellite interlink modules. Nevertheless, in terms of overall power requirement, this addition is not expected to change the total power requirement of the Communications plus DHS modules in each satellite.

128 Page 126 of 231 Table 12-3: List of equipments connected to the bus For the baseline configuration with two satellites, the power consumption of the subsystems has been computed for all the identified mission phases (See Table 12-4). Both satellite buses have an average power level in the range of W and the number of output lines are expected to be in the same range of magnitude. As for other equipment of a satellite platform, a recurrent approach on the power subsystems of the two (or three) satellites is expected to be advantageous in term of development time and cost.

129 Page 127 of 231 Table 12-4: Power Requirements on the bus Design methodology The vast majority of the operational orbits are envisaged to be dedicated to Sun-corona observations. Moreover, the absolute positions and pointing of the satellites during the FF demonstration manoeuvres is, at this design stage, not yet completely defined and therefore the Sun direction is unknown. A conservative and reliable approach would be to assume that the satellites might be in any attitude during the 20 hours long manoeuvre period. The main drawback of such an approach is an obvious over-design of the power subsystem, especially of the solar panels. Since the manoeuvres are expected to be limited to only 12 days, the manoeuvres can be adjusted to guarantee a minimum sun illumination of the solar panels during the 20 hours. Thus, the power subsystem will be sized according to the methodology illustrated in Figure To summarize, a first power architecture sizing (named the Optimized Design ) is made with only relaxed FF requirements (assuming that the SA will, in average, be kept facing the Sun with a maximum angle of 20 ). As a second step, the design is then increased to the maximum possible extension without involving any major configuration changes (named the Maximized Design or Baseline Design ). This final power subsystem configuration is then finally studied on a typical FF orbit. The outcome of this analysis will be the Sun attitude limitations to take into account for the FF manoeuvres.

130 Page 128 of Assumptions and Trade-Offs Figure 12-5: Power Design Methodology In line with the launch target and the cost constraints, use of existing and space qualified technologies shall be preferred. Hence, solar cells for power generation and battery cells for energy storage should be used. Li-ion battery modules with performances in-line with current off-the-shelf battery cells performances (18650HC from AEA and VES 100/140 from SAFT) are selected. AsGa TJ 27% AM0 are the space qualified cells having the best energy conversion and are therefore selected, on account of the major accommodation issues in Power architecture study A preliminary trade-off justifying the selection of a 28V regulated bus as oppose to an unregulated bus topology has been performed. The trade-off is based on qualitative factors differing between the two architecture types, but it also relies on accurate electrical simulations computed by the Powercap tool. This allows the performance expectations to be compared properly. For this preliminary comparison, the behaviour of an S3R-regulated architecture will be balanced against an S3R-unregulated (See Figure 12-6).

131 Page 129 of 231 Figure 12-6: The regulated and the unregulated voltage architecture considered options The two main drawbacks of a regulated architecture are typically: The BCR/BDR efficiency losses and their masses, resulting in a heavier and more costly PCDU as well as a slight increase in the battery capacity requirement. Nevertheless, these disadvantages are counterbalanced by the optimization of the dedicated DC/DC converters. In addition, three main rationales justifying this choice are listed here after: The selection of a regulated bus allows for an easier definition of the bus interface (voltage range, inrush current for example) for all the connected equipment, enabling an earlier start of their development phases Off-the-shelf equipment is easier to accommodate without modifications DC/DC converters integrated in the different units usually have poor performance compared to converters developed by power experts. As a result, an unregulated topology is even expected to require more energy from the batteries. The two main outcomes of the simulations with Powercap are: Assuming a full coverage of the Occulter s disc by solar cells, an unregulated architecture results in an average loss of 3 of the average angle constraint allowed in FF demonstration mode, compared to a regulated voltage design. Indeed, the 30 average Sun angle to take into account for the pointing of the disc toward the Sun during the FF manoeuvres decreases to 27. In the unregulated option, the total mass benefit of the battery and PCDU modules are expected to be lower than 2kg. Thus, the regulated bus topology has been selected for the Occulter in this preliminary study. For recurrent cost optimization, the same topology is also selected for the Coronagraph.

132 Page 130 of Two Satellites Solution (Baseline) Budgets and power system configurations In line with the sizing methodology defined in section , the Occulter and Coronagraph power subsystems are mainly designed according to the coronagraph science requirements. Due to the requirements related to the coronagraph observation mode, the solar cells of the Occulter can only be mounted on the circular disc itself. In fact, an additional deployable structure would increase the shadow area on the Coronagraph and bias the instrument measurements. In the optimized design (Table 12-5), the area available on the 1.2m diameter disc is compliant to the surface required for the solar cells. Including the packing factor of the mounting of the solar cells, 94% of the available surface is needed. In the maximized design, the shield is fully covered will solar cells and the critical average Sun attitude angle tolerated during the FF demonstration mode will be computed accordingly. During the coronagraph science observation phase, the Coronagraph satellite will receive a low solar illumination. Hence body mounted solar cells will be inefficient and deployable panels are preferred since they reach more illuminated areas. To facilitate the attitude control of the satellite, a configuration with two symmetrical and identical wings are selected. As a minimum, each wing shall be 55cm long (with a height of 75cm). This value is directly driven by the coronagraph science orbit and not by the 20 Sun angle of the FF demonstration orbits. According to the Coronagraph satellite dimensions, two stowed solar panels of 100cm can be mounted without impacting the overall configuration of the satellite. Therefore, two symmetrical and identical panels of 75x100cm are proposed for the baseline configuration (see Table 12-6). These additional solar cells will result in less restrictive Sun pointing requirements during the FF demonstration operations. The mounting of two SADM have been discarded due to the mass and volume penalties of such a system (Yoke, SADM, SADE and so on) and the limited benefit, since fixed deployable solar panels are already fulfilling the mission with a low constraint level on the satellite attitude during the FF manoeuvres.

133 Page 131 of 231 Table 12-5: Optimized Design: Power budget and design results The battery capacity is an outcome of the safe mode energy requirement. Including the redundancy cells and the capacity degradation, the battery module masses computed are 4.5kg for the Occulter and 3.8kg for the Coronagraph. The batteries are considered unchanged between the optimized and the baseline designs. Indeed, these batteries can supply the total required power in FF for more than two hours enabling manoeuvres of a couple of hours with the solar cells fully shadowed. Table 12-6: Baseline Design: Power budget and design results

134 Page 132 of Coronagraph satellite Solar Array (SA) Each panel is initially stowed against the body of the satellite. Once in orbit, they are immediately deployed. The RWE 27% space qualified cells (onboard Herschel Planck, Aeolus and Pleiades) are selected. No further qualification is required on the cell level for. On the other hand, the panels (and their hold-down and deployable mechanisms) are customized for the mission and a qualification process needs to be considered for these elements. Nevertheless, no major blocking point is foreseen Battery module Figure 12-7: Coronagraph SA configuration The Coronagraph battery module is an assembly of Li-ion battery cells. The type of cells and their arrangement in series and parallel has not been looked at in detail at this stage. Indeed, the existing small capacity cells allow the building of battery modules with performances (voltage range, capacity) close to the values required. Thanks to the MEX, Cryosat, Rosetta, VEX and other space programs, these cells are now fully qualified for the needs of, allowing the reuse of off-the-shelf cells and possibly also battery modules PCDU The PCDU includes the following: Solar array shunt sections (for up to 300/500W power capability) BCR/BDR Modules (200/300W maximum) Battery charging management unit Protections and switching capabilities of 20x28V regulated power lines (LCL + FCL) Internal power supply Internal TM/TC Control Unit The accommodation is derived from best existing space qualified hardware performances. A dedicated unit will need to be designed, customized for the mission. This module will require a qualification procedure and an EM should be considered.

135 Page 133 of Performance and simulations The mass breakdown of the power subsystem is provided in Table The main dimensions are displayed in the right-hand rows. Table 12-7: Equipment List Coronagraph Power Subsystem Depending on the eclipse duration, the maximum allowed average sun direction during the FF demonstration mode varyies from 67 (outside of the eclipse season) to 63 (after the longest eclipse). This loose requirement is expected to have a low impact on the FF manoeuvres. Figure 12-8: Coronagraph satellite: Coronagraph Science Orbit Simulation Results A coronagraph science orbit has been simulated in Powercap. The results are illustrated in Figure 12-8 with the following parameters displayed: Battery voltage Current battery

136 Page 134 of 231 Battery DoD Power generated by the SA Power required on the bus. The simulation starts at the beginning of the eclipse with a battery fully charged, and last exactly one orbit. The simulation of the coronagraph science mode highlights, that during the 12 hours when the Coronagraph satellite is partially shadowed by the Occulter, the battery is not in-use and remains fully charged Occulter satellite SA GaAs TJ RWE 27% solar cells are the selected cells for the Occulter satellite SA. The mass of the structure has been included in the configuration subsystem instead of the power subsystem. As a result, the mass computed for the SA only takes into account: the bare cells, their coverglass, the internal harness, the protection diodes, the interconnects and the cell adhesive Battery module Figure 12-9: Occulter Solar Array The same battery cell technology has been considered for the Occulter satellite energy storage module as for the Coronagraph satellite. The higher energy required during the safe mode implies a mass increase of 700g. The selection of the same battery cells in both satellites should induce some cost reduction. Further benefit can be achieved by re-using the full battery module in the two satellites. This advantage has to be counterbalanced with the 700g additional battery mass required in the Coronagraph satellite PCDU The same functions are implemented in the Occulter PCDU as in the Coronagraph PCDU. The lower bus power load and the smaller SA results in a PCDU slightly more compact and lighter (300g) than on the Coronagraph satellite. As for the battery modules, both PCDUs should have identical PCBs as far as possible Performance and simulations The Occulter power subsystem mass summary is provided in Table 12-8.

137 Page 135 of 231 Table 12-8: Equipment List Occulter Power subsystem Depending on the eclipse duration, the maximum allowed average Sun direction during the FF demonstration mode is varying from 34º (outside of the eclipse season) to 26º (after the longest eclipse). This stringent pointing requirement is expected to heavily influence the FF manoeuvres. The simulation of the FF orbit with Powercap, when the maximum tolerated average angle is reached during the FF manoeuvres, highlights that the battery is slightly discharging to compensate the lack of power generation from the solar panels. Nevertheless, at the end of the simulation, the battery is again fully recharged. Figure 12-10: Occulter satellite: FF demonstration orbit, Simulation Results with max average sun angle 12.4 Options For both satellites, the baseline power systems are the results of an optimization process minimizing cost, required development, maximizing use of off-the-shelf units, performances and reliability while limiting the attitude constraints on the FF manoeuvres. On the Occulter, the computed attitude constraints with respect to the Sun might be too strict and incompatible with the FF manoeuvres, while on the Coronagraph, these limitations are less stringent. If further analysis illustrates the incompatibility of the FF manoeuvres with respect to

138 Page 136 of 231 the Sun attitude requirements, a proposed solution is the implementation of body mounted cells on the other sides of the concerned satellite. For the Coronagraph satellite, a second alternative to increase the Sun angle capability is to place solar cells on the backside of the two solar panels. The main drawbacks of this option are the additional mass (for the solar cells but also for the harness) and also the solar cell raise the operating temperature, thus degrading the power conversion of the PVA. The third and most clever alternative for the Coronagraph satellite would be to tilt the two solar panels (Figure 12-11). The current power excess generated when sun pointed at 45% allows a panel inclination even greater than 45. This option has the advantage of not impacting the mass budget and improving drastically the FF capabilities. Nevertheless, this option is not considered as the baseline, since no need to improve manoeuvrability has been identified. Figure 12-11: Optional PVA configuration for Coronagraph satellite

139 Page 137 of STRUCTURES 13.1 Requirements and Design Drivers The structure shall fulfil the following requirements: Aim for simple load paths Withstand the design limit loads without failing or exhibiting permanent deformations that can endanger the mission objectives Ensure sufficient stiffness to decouple satellite modes from those of the launch vehicle. To be compatible with the VEGA and the Eurockot launchers, the stiffness requirements for the two stacked satellite (Coronagraph satellite and Occulter satellite) together with the Lisa Pathfinder Propulsion Module (LPF PM) are: o First axial frequency > 35 Hz. o First lateral frequency > 15 Hz. Provide support and containment for satellite units, equipment and subsystems Assumptions and Trade-Offs For a satellite, it is common is to apply a sandwich panel for the primary structure. A sandwich panel consists of two main elements. The outer parts of the panel are called face sheets, thin plates of metal or fibre reinforced plastics. The middle part of the plate is called the core, and consists of many strips of material that are combined to form a honeycomb. The two parts are combined by bonding the face sheets in such a way that the honeycomb core is closed. The core is normally honeycomb aluminium 3/ (density is kg/m 3 ). When local reinforcement is necessary, a higher density core is used e.g. 1/ (density of kg/m 3 ). For the face sheets, an aluminium alloy can be applied or a composite material. The major disadvantage of the aluminium alloy is its relatively high density and low stiffness, which imposes a mass penalty. Reasons to choose aluminium would be the following: Considerably cheaper Proven and easy manufacturing technology Better local load introduction Shorter manufacturing time Thermal stresses are minimised (in an overall aluminium structure) Supports thermal control (mounting of cold radiators) Sufficient material data available no (component) material development testing required Enhanced damage control by means of better repair possibilities Allows for simple implementation of local stiffening or strengthening elements at a very late stage to respond to shifts in the equipment accommodation. For the mass estimation of the structure in this study, a 20 mm high aluminium honeycomb core (3/ ) and Carbon Fiber Reinforced Plastic were chosen for the face sheets.

140 Page 138 of 231 Thickness of each face sheet is 0.5 mm. If mass is not a problem a sandwich with aluminium (e.g T6) face sheets, is a good alternative Two satellites solution (Baseline) Figure 13-1 show the overall launch configuration. Due to the launch stack of propulsion module together with two satellites, the Centre of Mass together with the stiffness of the individual launch element structures and the interfaces will determine the compliance with the launcher required stiffness. Coronagraph VEGA Launch Fairing Occulter satellite LPF PM Figure 13-1: Launch configuration The initial stiffness calculation for the two stacked satellites and the LPF PM results in a first lateral eigen frequency of 17.9 Hz. The assumption is made that the CFRP face sheet thickness of the LPF PM cylinder is 2 mm and that only the aluminium stringers, and not the CFRP panels, contribute to the stiffness of the two spacecraft. This shows that the preliminary design fulfils the requirement to be larger than 15 Hz. The axial frequency has not been assessed, since that involves a more detailed analysis requiring a more mature design state Coronagraph satellite The structure of the Coronagraph satellite will only carry its own sub-systems, and no other satellites during launch, and will have an interface with the Occulter satellite. The following two paragraphs explain the design, and provide a mass breakdown of the different parts of the structural design. The structural concept is similar to the Occulter satellite.

141 Page 139 of Structural design Figure 13-3 shows the primary structure of the Coronagraph satellite. Figure 13-2: Coronagraph satellite primary structure The structure is a skin-frame structure, with 6 CFRP sandwich panels, surrounding a skeletal framework made of aluminium corner stringers (members oriented in the vehicle s axial direction) and aluminium lateral frames. Two deployable solar array wings provide the necessary area for the required solar panel. The panels stabilize the framework and all equipment will be mounted on these panels. As described for the Occulter satellite, this concept provides easy accessibility to sub-systems, while maintaining the overall structural integrity. The mechanical interface with the Occulter satellite during launch is provided through four points at the bottom ends of the four corner stringers.

142 Page 140 of Mass breakdown Table 13-1 presents the mass breakdown of the Coronagraph satellite. Item Mass [kg] 6 panels stringers and 2 frames 9.1 Fasteners & inserts 4.8 Brackets 4.8 Total (including 10 % margin) 39.6 Table 13-1: Coronagraph satellite structure mass Occulter Satellite The structure of the Occulter satellite will carry the Coronagraph satellite during launch and provide an interface to the LPF PM. The LPF PM provides the interface to the launcher. The following two paragraphs explain the design, and provide a mass breakdown of the different parts of the structural design Structural design Figure 13-3 shows the primary structure of the Occulter satellite. Figure 13-3: Occulter satellite primary structure

143 Page 141 of 231 The structure is a skin-frame structure, with 6 CFRP sandwich panels, surrounding a skeletal framework made of aluminium corner stringers (members oriented in the vehicle s axial direction) and aluminium lateral frames. One of the side panels is a circular panel, the so called occulter disc, also providing the necessary area for the required solar panel. The panels stabilize the framework and all equipment will be mounted on these panels. During the AIT phase of the satellite, the panels could be removed for accessibility to sub-systems, while maintaining the overall structural integrity. The interface with LPF PM is realised using an adapter bolted to the bottom ends of the four corner stringers of the Occulter satellite. The mechanical interface with the Coronagraph satellite is provided through four points at the top ends of the same stringers Mass breakdown Table 13-2 presents the mass breakdown of the Occulter satellite. Item Mass [kg] 6 panels stringers and 2 frames 9.1 Fasteners & inserts 4.8 Brackets 4.8 Total (including 10 % margin) 39.6 Table 13-2: Occulter satellite structure mass

144 Page 142 of 231 This Page Intentionally Blank

145 Page 143 of THERMAL 14.1 Requirements and Design Drivers The Thermal Control Subsystem (TCS) shall guarantee that the satellites and all subsystem units and payloads work within their temperature ranges in every phase of the mission. The coronagraph payload requirements are the main design driver for the Coronagraph satellite: Lens to be kept constantly at around 20 C CCD box to be kept constantly at -70 C (3W power dissipation assumed). The power dissipated by the units is in turn dissipated through radiative surfaces having temperature ranges of -10 to 40 C (values assumed for both Coronagraph and Occulter satellites) and -70 C constant value for the Coronagraph CCD box Assumptions A standard semi-passive design is realized with Multi-Layer Insulation (MLI) blankets (different kind depending on the location on the satellites, see Figure 14-1), thermofoil heaters/sensors, black and white paints, washers, fillers and doublers The external thermal environment as seen in Figure 14-2 to Figure 14-4, as for a 24h orbit with: o Apogee altitude of km o Perigee altitude of 800 km o Inclination of 60 The satellites are in the orbit plane with their y-axis Sun pointing and the +Z axis oriented in the anti-earth direction The satellites are placed at an ISD of 120m Same design possible for the Occulter and the 3 rd satellite.

146 Page 144 of 231 Figure 14-1: MLI blankets composition Figure 14-2: External environment acting on the Occulter and 3 rd satellite radiators (+Z surface)

147 Page 145 of 231 Figure 14-3: External environment acting on the Coronagraph radiator (+Z surface) Figure 14-4: External environment acting on the Coronagraph radiator (-X surface)

148 Page 146 of Two Satellites Solution (Baseline) Coronagraph satellite Below follows a description of the thermal design of the Coronagraph satellite: White painted radiator of 0.35m 2 on the -X surface MLI (type 1) on the external surface not used as a radiator High temperature MLI bags (type 2) around the thruster boxes MLI tents (type 4) on the rechargeable batteries Black paint on internal surfaces and dissipative units White paint on the antennas and rear side of the solar arrays Heater/sensor lines for thermal regulation. The CCD needs a dedicated radiator. Its temperature needs to be kept constantly at -70 C. This white painted radiator will have a surface of 0.08m 2 and will be placed on the +Z surface of the satellite. The unit will be enclosed in an MLI tent (type 3), the end decoupled from the rest of the satellite. Heater/sensor lines will be used for thermal regulation. The mass budget can be found in Figure Figure 14-5: Coronagraph mass budget For the lens of the Coronagraph which has to stay constantly at 20 C, two steady state simulations have been performed. It has been assumed that the lens is directly connected to the satellite structure on the +Y surface. The structure is covered with MLI having an emissivity (ε) of 0.68 and an absorbtivity (α) of The lens is directly affected by the external environment. It has been assumed the lens has a value of 0.05 for α and a value of 0.8 for ε. In a hot case environment, a value of 4.7W of heating power needs to be applied on the structure in proximity of the lens, while a value of 5.7W has to be provided in a cold case environment. It has been estimated that a maximum heating power of about 11.88W is required (6.18W for the Coronagraph CCD unit box and 5.7W for the lens). The average of the power dissipated during the different modes allows the other units to stay within their temperature range.

149 Page 147 of Occulter satellite Below follows a description of the thermal design of the Occulter (the 3 rd satellite has the same design): White painted radiator of 0.48m 2 on the +Z surface MLI (type 1) on the external surface not used as a radiator High temperature MLI bags (type 2) around the thruster boxes MLI tents (type 4) on the rechargeable batteries Black paint on internal surfaces and dissipative units White paint on the antennas Heater/sensor lines for thermal regulation. The maximum heating power will be W. The mass budget can be found in Figure Figure 14-6: Occulter and third satellite mass budget

150 Page 148 of 231 This Page Intentionally Blank

151 Page 149 of TELECOMMUNICATIONS 15.1 Requirements and Design Drivers TT&C communication during all mission phases, in any mode and with almost any attitude is required Design should be kept as simple as possible in order to maximize the mission duration and reduce cost Data rates shall be optimised by giving realistic assumptions of on-board equipment and ground segment availability RF metrology is required for relative position calculation Bi-directional data transmission between the two satellites is required, with omnidirectional coverage Earth-pointing is not guaranteed during the whole mission. On the other hand the two satellites will be Sun-pointing during the whole Sun-corona observation phases and no fixed attitude can be guaranteed during the FF manoeuvres The orbit is a 24h HEO orbit: the maximum distance to the Earth is reached in the apogee and amounts to km Assumptions and Trade-Offs Data transmission assumptions During the FF demonstration manoeuvres, the generated telemetry is: o Optical metrology (at 10 Hz): 120kbps o RF metrology (at 1 Hz): 4kbps During the coronagraph science phase, considering that: o One composite image (1 polychromatic plus 4 monochromatic images): 15.5Mbit o During a coronagraph science orbit 12 hours are available for Sun observation. As baseline it is envisaged to keep the coronagraph on for a maximum of 6h per dedicated orbit. o One image is taken every 5 minutes during the coronagraph on-periods. In the worse case approximately 2 232Mbit is generated during one full coronagraph science orbit. The maximum distance to the Earth is km. All transmission data rates are calculated for that distance Antenna trade-off Considering: The two satellites will be Sun-pointing during the coronagraph science phase The attitude of both spacecraft during the formation flying manoeuvres is unknown, and not guaranteed to be Earth-pointing Steering mechanisms should be avoided in order to reduce system complexity, risk and mass.

152 Page 150 of 231 Traded-off antennas: Dish: discarded because of the necessity of a pointing steering mechanism due to their low beam-width (in the order of a few degrees) Helix: discarded because of their big size, weight, and difficulties in accommodation compared to patch antennas Patch: small, light and easy to accommodate. The main drawback is the low gain (with respect to helix antennas), close to 90 from boresight. The selected antenna is an LGA patch Frequency band selection For near-earth missions, the allocated frequency band for spacecraft operation is the S-band ( MHz for Earth-to-Space links and MHz for Space-to-Earth links communication). However, this S-band (which is shared by Space Research (SR) Category A, Space Operation (SO) and Earth Observation Services, plus high density terrestrial mobile systems) is currently highly congested, and sharing difficulties with fixed systems have already appeared. As a result, the ITU employs complex frequency coordination procedures in this band. Additionally, the Space Frequency Coordination Group resolves that whenever practical, bandwidths should not exceed 6 MHz, to reduce future congestion in the band. An alternative for the crowded S-band would be the X-band. Considering X-band versus S-band, the most favourable band of operation depends on the kind of antenna used at both ends of the link (ground and space). In this case, assuming constant aperture at the ground station and an LGA on board, the communications performances of S- and X-bands are almost similar in clear sky conditions (atmospheric absorption and rain losses are higher in the X-band). Although the S-band is quite congested, the requirement for telemetry downlink is quite relaxed so there is no real need to move to X-band. Staying in the S-band also guarantees the re-use of a lot of the existing flight hardware (for example from PROBA2). In conclusion, the S-band has been selected for both uplink and downlink. Note that since there are two satellites, in principle two frequency allocations will have to be requested. Considering the long passes and large coverage areas (due to the HEO), it is likely that these frequency allocations will be allocated solely to the mission and not to be able to be re-used by other LEOs Two Satellites Solution (Baseline) TT&C architecture In the two satellites solution, the TT&C architecture is based on the following points: Fully redundant TT&C capacity on each satellite: this includes two receivers in hot redundancy and two transmitters in cold redundancy. The choice of full redundancy on each of the two satellites is based on: o The requirements during (emergency) LEOP o The benefit of having additional downlink capability. Medium-rate inter-satellite communication: this functionality is already included in the RF metrology system with a data-rate in the order of a few kbps.

153 Page 151 of 231 The figure below shows the TT&C concept for the two satellites. Figure 15-1: TT&C concept for two satellites On-board TT&C system The on-board TT&C system for the Occulter and the Coronagraph is basically the same: both carry full TT&C capabilities, but with a different frequency allocation. The on-board TT&C system is based on: 2 redundant transceivers 2x2 hemispherical LGAs for full omni-directional coverage Navigation is performed by: o GPS receiver (when below constellation) o Range rate based on the Zelinda algorithm (experiment on Smart-1) Modulations and coding The used modulations have been chosen from the ECSS standard, RD[39], considering that this is a CCSDS category-a mission. The used modulations are: Uplink: NRZ/PSK/PM Downlink SRRC-OQPSK with roll-off 0.5. For downlink telemetry, the coding scheme selected is the CCSDS standard Convolutional-Reed Solomon concatenated code with interleaving depth of I = 5 (see RD[40]). This coding scheme is compatible with the ESA ground station Network and guarantees a Frame Error Rate of 10-5 at Eb/N0 = 2.8 db.

154 Page 152 of 231 Uplink Downlink Modulation PCM/PSK/PM SRRC-OQPSK Forward Error Coding - (CC(1/2,7),RS(255,223)) with I = 5 FER BER 10-5 FER 10-5 Synchronisation ASM ASM Table 15-1: Up-and downlink modulation and coding On-board antenna selection and location Surrey Satellite Technology Ltd. SPA-Series S-band patch antennas are considered as baseline, RD[52]. The mass of this antenna is 80g and its dimensions are 82x82x20mm. Figure 15-2: S-band patch antenna from SSTL The gain pattern of the antenna is shown below: 7 dbi gain at boresight 2 dbi gain at +/ dbi gain at +/- 90. Figure 15-3: Gain pattern of S-band patch The requirements on the TT&C antenna location are to provide full duplex communications during all mission phases, modes and with any attitude. As the two satellites will be Sun-pointing during most of the mission time, one has to take care that (as far as possible) a downlink to Earth at nominal rate will always be available. Since the two satellites are not Earth pointing, this means that in case only two LGAs are mounted (for example one on the satellite +X-axis and one on the X-axis), the Earth can be at 90 with

155 Page 153 of 231 respect to the antenna boresight. Designing a TT&C system based on this worst case scenario (that is with a very negative antenna gain) would lead to an over-design of the system. Instead, to ensure nominal downlink capabilities during every phase of the mission, a multiple antenna arrangement with two pairs of LGAs is proposed: One Tx LGA pair transmits along the X-axis of the satellite The second Tx LGA transmits along the Y-axis of the satellite To limit transmission chain losses, switches are used to select the proper LGA pair based on the position of the satellite with respect to the Sun and to the Earth. Using this multiple antenna arrangement, the maximum angle at which the Earth is seen from any of the available transmitting antennas is 45 with respect to the boresight. Taking this into account and allowing for some margin, the link budget for nominal downlink operations will be calculated based on 60 off-pointing angle, corresponding to an antenna gain of 2dBic On-board S-band transceiver Due to the low-cost requirement for this mission, the S-band transceiver proposed for this mission is an upgrade from the PROBA2 one. The PROBA2 TM/TC unit contains: 2 redundant S-band transceivers a Radio Frequency Distribution Unit (RFDU) Sseparate DC-DC converters for each transceiver. The figure below shows a functional block diagram of the S-band transceiver. Figure 15-4: Functional block diagram of S-band transceiver (from PROBA2)

156 Page 154 of 231 The total mass of the PROBA2 TM/TC unit is 5.5kg. This excludes the four low gain antennas, but includes the large RFDU. The power consumption is: 23W in transmit mode (5W Tx power at amplifier output) 3W in hot redundancy receive mode. Finally, Figure 15-5 shows the full TM/TC unit. In the bottom figure, the two upper boards comprise the redundant S-band transceivers (with power supply connectors PL1 and PL2 and Data/Monitors connectors SK1 and SK3). The bottom board houses the RFDU and feeds the four antenna ports SK5, SK5, SK7 and SK Link budgets Figure 15-5: PROBA2 TM/TC unit Link budget during nominal operations The details for the down- and uplink under nominal conditions can be found in Table References RD[39] to RD[46] will define the link. Nominal conditions include: 15m Perth ground station Link distance: km LGA gain = 2dBic, both in transmit and receive (relevant to 60 degrees off pointing)

157 Page 155 of 231 High-power mode: Tx power = 6W 1 Downlink: 250kbps Uplink: 64kbps. Hence the downlink is designed to provide a data-rate of at least 250kbps at maximum link distance with a 60 off-pointing angle. This allows for: Real-time downlink of telemetry generated by: o optical metrology at 10 Hz (120 kbps) o RF metrology at 1 Hz (4 kbps) Total coronagraph science return of 5400 Mbit during two 3-hour data dump passes over the ground station. Concerning the uplink, 64 kbps is available during the ground station passes which is considered sufficient to regularly upload FF scenarios to the two satellites. Since the two satellites are in the same orbital slot, it is possible to command both simultaneously. Assuming that both satellites operate at the same receiving frequency, a multiple access strategy can be implemented to command each satellite individually. Possible options for multiple access are: Discrimination based satellite ID a TDM approach. Uplink Downlink Frequency 2110 MHz 2290 MHz Spacecraft LGA gain 2 dbi 2 dbi Tx power 398 W 6 W Coding No coding Concatenated CC(1/2,7),RS(255,223) with I = 5 FER Bit rate (nominal operation) 64 Kbps 250 kbps Link margin 8 db 6 db Table 15-2: Characteristics of the S-band link during nominal operations Link budget during low-rate and contingency operations The details for the down- and uplink under contingency conditions can be found in Table During contingency, it is assumed that: 15 m Perth ground station Link distance = km LGA gain = -3dBi (worst-case both in transmit and receive) Low-power mode: Satellite Tx power = 2.5W Downlink rate: 8kbps Uplink rate: 4kbps. 1 Power at the output of the amplifier, transmission chain losses not included.

158 Page 156 of 231 Uplink Downlink Frequency 2110 MHz 2290 MHz Spacecraft LGA gain -3 dbi -3 dbi Tx power 398 W 2.5 W Coding No coding Concatenated CC(1/2,7),RS(255,223) with I = 5 FER Bit rate (low rate) 4 Kbps 8 kbps Link margin 16 db 12 db Table 15-3: Characteristics of the S-band link during contingency operations RF metrology Next to relative navigation, the RF metrology package also provides medium rate-data transfer capabilities between the two satellites, in the order of a few kbps. For more details refer to the RF metrology chapter Options The use of Turbo Codes ½ for TM would allow a theoretical data rate increase of about a factor 1.4, by reducing required telemetry data Eb/N0 by1.6 db. This code is not included in the baseline because: It would add complexity to the satellite Turbo codes are currently not supported by the ESA ground station network. The current data-rate available seems sufficient List of Equipment In Figure 15-6, a summary of the communications equipment (including the RF FF subsystem and the GPS receiver) with their masses can be seen. Total mass with margin is 14.4kg; in addition harness mass is estimated to be 2kg. Figure 15-6: Communications and RF metrology equipment.

159 Page 157 of Three Satellites Option The main proposed change for the three satellites solution is to include a high speed intersatellite link between all the satellites within the formation. By providing high-speed data transfer capabilities between the satellites, the redundancy within the formation can be made use of to reduce some of the redundancy in each satellite. In this scenario, the inter-satellite link (ISL) capabilities will allow fast transfer of commands and data between all of the satellites in the formation. Given these ISL capabilities, there is no need for a redundant S-band transceiver on-board each satellite, as its function can be taken over by the inter-satellite transceiver. In case the link with the Earth fails on one of the satellites, due to a malfunction in the on-board S-band system, a back-up link can be provided via each of the other two satellites using the ISL. Moreover, even if two out of three satellites loose the capability to communicate directly with ground, full control over each of the satellites in the formation can be kept. The figure below shows the TT&C concept for the three satellites solution. Figure 15-7: TT&C concept for three satellites solution In conclusion, in the three satellites solution, the proposed design per satellite includes: One S-band transceiver providing direct two-way communication with Earth One inter-satellite transceiver providing high speed, full duplex communication capability with the rest of the formation. The main advantages of this approach are: Mass saving: the redundant transceiver providing communication with Earth is replaced by a much smaller and less power consuming ISL transceiver.

160 Page 158 of 231 Higher reliability: for larger formations, the fault-tolerance of the total formation increases using this approach. For example, assuming three satellites in the constellation, the result is a 2 fault-tolerant communications system for the entire formation with respect to a 1 fault-tolerant system assuming the classical approach (full redundancy per satellite). The design of the ISL can be based on terrestrial network protocols (for example WiFi or 802.xx). Various research activities are already on-going in this field, with the purpose of reusing these terrestrial network protocols for space applications. For an overview of the activities within ESA, see RD[56]. More info is also provided in the Data Handling System Chapter. The equipment list for the three satellite solution is shown below (this configuration is valid for all three satellites). Figure 15-8: Equipment list three satellite solution

161 Page 159 of DATA HANDLING SYSTEM 16.1 Requirements and Design Drivers For the development of the DHS architecture, the driving requirements that have been considered are mass and cost reduction. The performances requirements (in terms of data throughput, radiation, harness and system complexity) are quite standard and allow reuse of existing components and building blocks. It shall be noted that an intelligent design of the DHS will allow further cost savings in AIT and operations. These numbers are difficult to estimate at this stage, but possible optimization should be kept in mind from the beginning of the design process. is also a test opportunity for validation of promising technologies and/or concepts. There can be additional risks and costs coming from that, but missions like are sometimes the only occasion for new and/or experimental hardware to fly Assumptions and Trade-Offs Thanks to the short duration of the mission, the radiation environment is quite benign, as can be seen in Figure A shielding thickness between 2 and 4 mm Al is enough to have a total dose (1 yr) less than 10 KRad. Figure 16-1: 4pi total dose for orbit profile

162 Page 160 of 231 It shall be considered that with less than 10 krad, use of COTS components and systems is possible, and shall be traded off from the beginning of phase A. To have an example of the possible gain, the same performances in terms of computing power can be obtained with a COTS SBC with a factor of 5 less mass and power with respect to space qualified hardware with QML-Q components. The use of COTS would also allow an easier implementation of integrated architectures (centralized power distribution, TMR, functional redundancy), and could be of great benefit for exploring spin in from commercial technologies. Among the terrestrial technologies that are already proven, the CAN Bus provides some key advantages with respect to more traditional buses such as MIL-STD-1553B. It provides a similar performance, but at a very low cost, and an improved robustness due to its multiple error detection mechanisms. Thanks to the wide use of CAN Bus in terrestrial applications, numerous development and analysis tools are available that support and reduce the cost of all the engineering process, from specifications to AIT. Currently, an ECSS standardization is in progress for making the best use of CAN in Space (see website). For the DHS, the basic assumption is that an identical architecture can be used for the two satellites. This (together with the maximization of the reuse of existing HW, if space qualified HW is chosen) will allow a development process that will minimize the non recurring costs (EM development, test, basic software and so on) Two Satellites Solution (Baseline) The model for the OBC can be the GSTB-V2B (also known as GIOVE-B) ICDU. Both the size and the performances of the unit will fit needs (downgrading the I/O needs) without any new development. Then the proposed DHS units can be: An Integrated Control & Data Management Unit, that includes: 2 x Processor Module (ERC32-25MHz) 1 x TTFG RM SGM 4 x HPCG/IOB I/O modules Mass memory (8/16 Gbit, EDAC protected, internally redundant). A fully redundant box, internally X-strapped via spacewire/1553 is 13 kg, 35 W, including a dedicated DC-DC converter module.

163 Page 161 of 231 Figure 16-2: Block scheme of the proposed architecture. The SSMM can be integrated into the ICDU. This choice saves mass & power, increase slightly the failure probability and reduces the number of possible manufacturers of the equipment Three Satellites Option From a DHS point of view, the three satellite option is much more intriguing. Figure 16-3 shows how a formation of non-redundant identical satellites, thanks to fast ISL and file transport protocols (like CFDP) could be managed from ground as a single entity. Figure 16-3: Operational scheme for a network of three identical satellites

164 Page 162 of 231 The architecture for each satellite is much simpler, no differences, no redundancies, same software, with the possibility to share resources (like memory and bandwidth) between the three satellites. Figure 16-4: Block scheme of the DHS for a network of three identical satellites. A fast comparison is possible between the two designs: the classical one (as for the two satellites option) and the distributed one with ISL(as suggested for the three satellites option). The classical design requires the following hardware: Processor x 2 TTFG x 2 RM/SGM x 1 IO boards x 4 X-Strap Backplane SSMM at its EOL capacity (2x) MM Control Module. which results in a total of 13 Boards with a CPU box mass of 14kg (per satellite). The distributed design on the other hand requires: Simplex CDMU Processor SGM (1/2) SSMM (1/3 capacity) IOB x 2

165 Page 163 of 231 ISL (1/2) TTFG (only one!) moved inside the SBTA Two patch antennae (4pi). Thus it is estimated to have 6 Boards/3 Boxes + 2 antennae for a Total Mass of less than 7kg per satellite with a harness being approximately 70% lighter. However, these figures do not take into account all other savings (Simpler SW, Simpler Harness, cheaper integration, test & development). The biggest gain is in reliability. It can be demonstrated that the distributed system (the system being the set of the three satellites) is 2 x Fault Tolerant.

166 Page 164 of 231 This Page Intentionally Blank

167 Page 165 of RISK 17.1 General Approach Two main topics have been dealt with in the Risk discipline in the frame of the Mission: 1) The definition of the Safety and the Mission Success Criteria together with a preliminary set of requirements to achieve them. 2) A preliminary approach for the implementation of Risk related design criteria in the project/mission Definition of the Safety and of the Mission Success Criteria Safety The main Safety issues are summarised below: Safety can be jeopardized by hazardous events or failures whose consequences can lead to the following events: S1) The potential to cause loss of life, permanent disabling injury to personnel or occupational illness, loss of launcher or of launch site facilities, severe detrimental environmental effects S2) Temporarily but not life threatening injury or illness, major damage to public or private properties, major detrimental environmental effect Mission success criteria For a proper approach for the identification of the Mission Success Criteria the following key issues have been considered: M1) Number of times the Formation has to be composed during the 19 months mission M2) How long each FF has to last to be considered successful M3) The precision that has to be required for the FF M4) The definition of the minimum set of manoeuvres to be performed to successfully validate the FF manoeuvres M5) Minimum set of experiments to be performed during the mission. For each Mission Success Criteria identified a dedicated requirement has been identified and proposed. It has to be noted that these requirements are to be considered as guidelines, the values referenced are provided for better clarity but will have to be defined in a second phase, when the details of the mission will be defined Requirements Proposed for Mission Success Criteria, M1 to M5 M1) Number of times the Formation of two (or three) satellites has to be composed during the 19 months mission. Example of Requirement to implement criteria M1):

168 Page 166 of 231 The Formation of the Satellites (FF) has to be composed a minimum of 12 times during the 19 months of the mission with a minimum of 2 times per month. M2) Duration of each FF (How long each FF has to last to be considered successful). Example of Requirement to implement criteria M2): The design of the satellites will ensure the FF operations and will guarantee for each FF a minimum duration of TBD hours (for example 12 hours). M3) Precision required for the FF. Example of Requirement to implement criteria M3): The relative position accuracy of the satellites during the FF shall be: Relative Position Accuracy: min =.. mm Max=.. mm Relative Attitude Accuracy: min =.deg. Max=..deg. M4) Definition of the minimum set of manoeuvres to be performed to validate successfully the FF. Example of Requirement to implement criteria M4): In order to validate the capabilities of the FF and to define it successfully, the mission and satellite design shall guarantee a TBD number of manoeuvres. M5) Minimum set of experiments to be performed. Example of Requirement to implement criteria M5): design and FF operations shall be optimised to guarantee high science return from the coronagraph instrument. A list of experiments and their relevant level of importance can be defined Identification of Undesirable Events and Relevant Severity Categories The next step, after the identification of the Mission Success Criteria and requirements, is to identify the Undesirable events that will jeopardize them and to associate a severity category to each Undesirable Event. The following Undesirable Events shall be defined as Mission Critical Category 2 Events UnEv1) All those events or failures whose occurrence will lead to the total loss of one of the satellites. UnEv 2) All those events that will lead to inability to perform a minimum number of FF manoeuvres and maintain the formation for a minimum TBD period. The following Undesirable Events shall be defined as: Mission Major Events Category 3 Events, Mission Degradation

169 Page 167 of 231 UnEv3) All those events or failures that will lead to the inability to maintain the formation or to perform the FF manoeuvres, for the minimum required time and with the required precision. UnEv 4) All those events or failures that will lead to the inability to perform the minimum set of scientific experiments The approach for mitigating the Undesirable Events occurrence For Mission Critical Events Category 2, a combination of redundancy implementation (single failure tolerance approach) together with the selection of components (reliable components), back-up solutions (including recovery procedures) has to be implemented to optimize the design within the mass and resources available. For Mission Major Events Mission Degradation - Category 3: No Failure Tolerant (FT) design has been required, but the design shall anyway address all those solutions that, compatible with the state-of-the-art actually in use in ESA space projects, will guarantee the minimization of the occurrence of all the identified failures leading to Mission Degradation. Remark: In addition, every time the mass and budget will allow it, a redundancy approach is suggested in order to avoid the occurrence of Mission Degradation Events Preliminary Approach for the Implementation of Risk Related Design Criteria Two satellites solution (baseline) For this configuration, the mass margins available are enough to allow a classical approach in accordance with the previously identified Safety and Mission Success requirements, that is: 1 FT design (and Fail Safe/Safe Life design for the Structural and Mechanical parts) for the subsystems whose failures can cause the Mission Critical Category 2 - Undesirable Events. For the Safety related Requirements: All those events or failures that will lead to consequences listed in S1) shall be categorised as Catastrophic - Cat I and be protected by a two failure tolerant design or three safety barriers All those events or failures that will lead to consequences as listed in S2) shall be categorised as safety Critical - Category II and protected by a 1 failure tolerant design or two safety barriers Structures and Mechanisms shall be designed to cope with design for minimum risk criteria (safe-life/fail-safe design).

170 Page 168 of Three satellites option As this option puts more heavy constraints in terms of mass and budget, the risk mitigation criteria have to be carefully implemented to maintain a sound approach, trying not to kill the mission by posing too heavy constraints that will result in non-feasibility. A minimum and smart use of redundancies is needed to optimize the cost (in terms of mass) benefit (in terms of increased reliability) approach. Therefore the following approach is suggested: 1) To implement a Data Handling System at satellites level. This means, three identical and not internally redundant DHLs, one per satellite. Each satellite s DHL is alone able to manage the complete formation (one of the three is seen as Master and performs the GNC functions plus experiments for the complete formation). In case of Master DHL failure, the DHL of one of the other two Satellites will take the control of the formation 2) Satellite interlink that allows complete communication and data exchange between the satellites, to allow control from each satellite 3) All the satellite equipments and payloads should be connected to the satellite internal bus and this internal bus in turn should be connected to the satellite intercommunication link. This solution will allow managing one satellite in case of failure of its DHL, and will allow communicating with it from ground by transponding with one of the other two. 4) A further possibility to enhance the mass reduction: If at least two of the three satellites can be designed and built identical, they could even be proposed as the redundant one of the other: a two satellites system with a third one seen as redundant with its twin. This means that the nominal formation is intended with two satellites (one main satellite with some internal redundancies plus two other satellites identical and redundant of each other). All the experiments and manoeuvres are performed with three satellites but in case of failure in one of the twins, the mission is not considered lost. The advantage is an additional mass saving, as the two twins have no internal redundancies. This concept could open the possibility for future FFs with multiple satellites to implement a zero failure tolerant internal design by using one of the satellites as a redundancy for the group.

171 Page 169 of SIMULATION AND VISUALISATION 18.1 Introduction A prototype Functional Engineering Simulator (FES) has been developed in the framework of the assessment study. The simulator is used in this study for the investigation of ground-station visibilities versus orbits, the calculation of the solar phase angles and the visualization of the relative motion of the different satellites during the creation and maintenance of the formation in orbit. The FES mainly consists of the functional models for calculating the position and attitudes of two or three satellites, and models of some subsystems. Moreover, an optical experiment was added to account for the analysis of data acquisition during coronagraph science observations. The solar system was modelled in a light way, consisting of the positions and attitudes of the Earth, Moon and Sun, and the ground-stations under investigation. Simple graphical models of the foreseen satellites (both the Coronagraph satellite and the one or two Occulter satellites) have been implemented in the 3D visualisation. The 3D visualisation is used to illustrate the creation of the formation for the coronagraph science observation phase in a more intuitive way Requirements on the FES During the study, the FES has been mainly used for performing the following tasks: investigation of ground-station visibilities for different orbits, and calculation of the solar aspect angles in orbit. In order to do so the following characteristics were studied: 1. Ground-station visibility versus the position of the satellite in orbit. This was done for the Kiruna, Perth and the VillaFranca ground-stations, taking into account their characteristics such as masking angle 2. Moreover, during the visibility periods the elevation angle of the satellite as seen from the respective ground-station was determined. 3. For the reference orbits the Phase Angle was calculated, being the angle between the vectors Earth to Sun and the vector Earth to Satellite. The following sections describe the implementation in the FES and the results of the simulation runs Description of the FES This mission and system simulator has been built using the SimVis software package to assemble the components of the simulator, and uses Simsat as the simulation kernel. The FES mainly consists of the environmental models for two or three satellites, containing functional models of the orbits and attitudes of the satellites. These include perturbed Keplerian orbit propagators, and attitude determination in a kinematical way. The satellite subsystems modelled consist of a Data-management subsystem, a Communication subsystem (including steerable antenna, receivers and transmitters), a Power subsystem (including solar arrays and batteries), and a Propulsion subsystem. Moreover, an optical experiment was added to account for the data acquisition analysis. The solar system was modelled in a light way, consisting of

172 Page 170 of 231 Earth, Moon and Sun attitudes and positions, and included three ground-stations: Kiruna, Perth and VillaFranca Simulation Results In this section some of the simulation results are shown. The used reference orbit for the two or three satellites formation is the following: Perigee distance = km Apogee distance = 77150km Inclination = 63º Argument of Perigee = 270º (90º for Perth) (Initial) True Anomaly = 0º Reference Epoch = 21/06/ :00: Ground station visibility Figure 18-1 shows the visibility of Kiruna from the reference orbit. In the plot, the orbit track is shown in black when the ground-station is not visible and a blue line is shown when it is. Due to the nature of the orbit and its period the visibility stays the same over the year. The yellow circle indicates the position of the Sun at that date.

173 Page 171 of 231 Figure 18-1: Kiruna ground-station visibility as a function of the satellite position in the orbit. Blue line means visible, black line not visible From Figure 18-1, it can be seen that for a long period of the orbit, including the apogee passes, the Kiruna ground-station is visible. Figure 18-2 shows the elevation angle for one orbit, on the 21 st June 2010.

174 Page 172 of 231 Figure 18-2: Elevation angle of the satellite as seen from the Kiruna ground-station. Please note the masking angle of 5º The following figures show the results of the same analysis for the Villafranca and Perth groundstations. Figure 18-3 shows the visibility of Villafranca from the reference orbit. Figure 18-3: Villafranca ground-station visibility as a function of the satellite position in orbit. Green line means visible, black line not visible. Figure 18-4 shows the elevation angle of a satellite as seen from the Villafranca ground-station on the 21 st June 2010.

175 Page 173 of 231 Figure 18-4: Elevation angle of the satellite as seen from the Villafranca ground-station. Please note the masking angle of 5 Figure 18-5 shows the visibility of the Perth ground station from the reference orbit and Figure 18-6 the elevation angle on the 21 st September For details, refer to the Mission Analysis Chapter. Figure 18-5: Visibility of the Perth ground-station as a function of the satellite position. Red line means visible, black line not visible.

176 Page 174 of 231 Figure 18-6: Elevation angle of the satellite as seen from the Perth ground-station D Visualisation By connecting the simulator to the OpenIGS visualisation tool, the graphical CAD-models, as obtained from the configuration chapter, can be dynamically visualised in a 3D graphical way. In the following images, the red-green-blue axes define the satellite body frames. The scales of the different satellite components are deliberately exaggerated to highlight their relative positions. Figure 18-7 shows the configuration of the two satellites during the perigee pass of each orbit. In the shown case, both satellites are in the same orbit with a difference in true anomaly. The Occulter satellite is trailing the Coronagraph satellite.

177 Page 175 of 231 Figure 18-7: Configuration of the Occulter (trailing) and the Coronagraph satellite during a perigee pass. Figure 18-8 shows the Creation of the Formation, where the Occulter satellite is positioned inbetween the Coronagraph satellite and the Sun. The white straight line shows the direction towards the Sun. Both yellow lines indicate the track of the satellites. Figure 18-8: Creation of the formation Concluding, Figure 18-9 shows the resulting configuration during the coronagraph science phase.

178 Page 176 of 231 Figure 18-9: Configuration of the formation during the coronagraph science phase

179 Page 177 of GROUND SEGMENT AND OPERATIONS The ground segment and operations infrastructure for the Flight Operations Segment (FOS) of the mission will be set up by ESOC, and will be based on the extension of the existing ground segment infrastructure, customised to meet the mission specific requirements Requirements and Design Drivers The ground segment shall guarantee safe operations during LEOP, commissioning and the routine phase. A single VEGA launch into a 300km LEO has been selected, with a LPF Propulsion Module (PM) used for orbit raising to a 800x70772km HEO. The satellite will then separate from the PM and immediately transition into coarse mode metrology operations. Both satellites will communicate with the Ground using redundant S-band transceivers and an LGA. The Inter-Satellite Link (ISL) has a capacity of a few kbps and is only foreseen as being used for the RF metrology function. The mission has the following characteristics: Launch in 2010 and a predicted lifetime of approximately 19 months: LEOP & Commissioning, 1½ years of coronagraph science 12hrs/day; 3 sets of 6 non-contiguous days of FF 20hrs/day LPF-type LEOP with progressive apogee raising burns using the LPF propulsion module. Minimum duration of 10 days The operational orbit is a 24h HEO optimised for maximum visibility from Perth. There will be a daily 20.4hrs telecommunications window The orbit will remain fixed in the inertial frame. Therefore, the local time of apogee for a fixed location on ground will progressively change throughout the year. Despite the long periods of daily visibility possible, the satellite will have to be highly autonomous due to the FF demands of the mission and the chosen orbit. Ground station(s) will be required for communications with the satellites, providing different periods of visibility according to the requirements of the different mission phases: LEOP and Commissioning FF tests Coronagraph Science observations. 6hrs/day is required for the dumping of the data from the scientific observations Assumptions and Trade-Offs The main assumptions considered for the design of the ground segment for the mission are the following: Satellite design and development is a low-cost, technology demonstration mission with a short duration. Nevertheless, its Ground Segment and Operations will still need development the same as any other longer

180 Page 178 of 231 duration mission. In addition, it will be the first of its kind flown by ESA so there will be limited inheritance possible from previous missions. There are, however, good practices that can be followed which are aimed at keeping operations preparation costs down: Strict adherence to ECSS defined standards in the design and development phases, to allow the exploitation of generic operations infrastructure and systems to the maximum extent possible The spacecraft database shall comply to the SCOS2000 MIB ICD Procedures shall be provided by the Prime in MOIS format Free and full exchange of mission developed software (for example simulator models) between all entities involved in the project Reuse of previous ESA technology (for example spacecraft bus) as much as possible. The beneficial implications of these steps will be assumed Operations implementation It is assumed that the mission is comparable with the LPF mission from a programmatics point of view Due to its importance for future Observatory/Survey and (to a lesser extent) Earth Observation missions it is assumed that the control team will be comprised of members from these mission families It is assumed that the core teams will be organisationally as close as possible/practical to the H/P and GAIA control teams It is assumed that a lot of the development work done for the operations of the Propulsion Module in the LPF mission can be applied to the mission LEOP It is assumed that the satellites will be launched in early 2010 A 19 months mission lifetime, includes LEOP, Commissioning, FF tests and routine, scientific observations The two satellites will be launched by a single VEGA launcher from Kourou Orbit raising of the satellites from the LEO reached by the VEGA launcher up to the operational HEO will take a minimum of 10 days. A LEOP of 12 days will be assumed so as to include margins in the timeline It is assumed that 3 ESTRACK antennas will be used to support the LEOP activities. The S-band stations will be selected to give suitable visibility when monitoring and executing the sequence of apogee raising burns No dedicated backup station will be considered for the routine, coronagraph science observation phase. Satellite emergency cases will be supported by the network as per priority rules HKTM and coronagraph science data acquisition from the Perth 15m antenna (S-band) is the ESOC baseline It is assumed that during any critical FF tests the NNO 35m antenna can be baselined as the backup station. This would depend on the usage of the station at that time and the scheduling of the tests would have to be done so as not to clash with other missions using NNO

181 Page 179 of 231 The satellite will have a 6h Daily TeleCommunication Period (DTCP) during the routine phase The ISL is not designed for the passing of telecommand data between the two satellites. It is, therefore, assumed that it will be possible to simultaneously achieve communications lock with both satellites using the same ground station. In this instance, it is also assumed that there will be sufficient capacity at the station for full TT&C with both satellites, as well as the dumping of the recorded data from the Coronagraph instrument within the 6 hours window The standard rate of 4kbps uplink is assumed. The implementation of modifications that would allow the maximum uplink rate to match the maximum possible receive rate by the Communications subsystem (64kbps) is an option It is assumed that TT&C links and high-rate data (250kbps) downlinks are possible during FF manoeuvres and coronagraph science operations It is assumed that the on-board mass memory can store at least two days of coronagraph science observations It is assumed that there will be a high level of on-board autonomy for all elements implying the following: o Feedback loops shorter than 24 hours are not required by any element during noncritical mission phases o On-board orbit determination and control based on GNSS measurements taken during the short and high-speed perigee passes o Automated entry into a safe orbit in the event of an on-board collision warning. The composition of the Flight Control Team during mission preparations and mission operations will be determined by the criticality of the operations and the possibilities of sharing the team with other missions such as LPF It is assumed that it will be possible to set-up the LEOP timeline such that critical operations can be covered by the main FCT (A-Team) It is assumed that it is sufficient that the LEOP Back-up Team (B-Team) is comprised of FCT members from another mission (such as LPF) with limited P3 training and that they will be involved mainly in monitoring activities SPACONs will be shared with the other missions in the Family of Observatory/Survey Missions Use of the SCOS2000 Multi-Mission, Multi-Domain Mission Control System (MCS) is assumed, the same MCS is assumed for both satellites. It is assumed that some automation will be available including: Routine Pass Operations and post-pass reporting capabilities Use of a simulator using the SIMSAT infrastructure (including emulation of the on-board processor) is assumed It is assumed that the on-board processor is either an ERC-32 or a LEON Hardware usage will be shared with H/P where possible (for example MPS, back-up system for the Data Dissemination System)

182 Page 180 of 231 The mission planning scenario will be divided into different levels, namely long, medium and short term planning. Months before each event (for example FF tests), a baseline planning will be established and this plan will be refined and prepared for uplink shortly before the event Mission planning will be supported during the normal working hours of the FCT Reaction times to anomalies during critical mission phases (for example LEOP) will be dependant on ground station coverage, to what extent the problems are detectable in the HKTM, and whether Flight Control/Contingency Recovery Procedures are already available It is assumed that the first set of 6 FF test days will be performed during the Commissioning phase as part of the commissioning activities of the FF subsystems and that the remaining set can be spread out throughout the routine coronagraph science phase It is assumed that the routine phase can be executed as an offline mission with no need to monitor the satellites in real time. This is only if the Commissioning and the first set of FF tests have gone well and the FF platform is operating properly Daytime-only operations are preferable for a low-cost mission. This means that passes outside of office working hours should be avoided, even monitoring passes as these would still require a SPACON on-console and an engineer on-call just in case there is an anomaly that has to be investigated. In the extreme case, this would mean no passes over the weekend. It is assumed that daytime-only operations during the working week can be applied to some, if not all, of the routine coronagraph science phase It is assumed that the FF test days taking place during the Commissioning Phase and the routine coronagraph science phase will require full-time, on-console engineering support for the whole 20h duration of the test LEOP recommendations The LPF-like LEOP of does not have the same high-level of risk as with the LPF mission itself: it is assumed that the VEGA launcher will have a much lower launch dispersion than ROCKOT, so there is less chance of time being lost due to search activities at first acquisition; communications are in S-band so more than just 3 ground stations will be available for TT&C per orbit and this, along with a launch from Kourou, it is assumed, will result in an acceptable duration before first acquisition. Nevertheless, there are still some factors that make this a riskier than normal LEOP: It is to be assessed if the low initial perigee altitude, despite the increased visibility possible due to the S-band communications, could mean that the majority of the Apogee Raising Burns (ARBs) to achieve HEO will have to be performed in the blind. Such a LEOP comes with a significant increase in the workload for the entire FCT, compared to a conventional LEOP. This increases the risk of error in the event of an anomaly. The following recommendations are therefore made: Add margin from the start in operations planning and on-board systems such as propulsion and power:

183 Page 181 of 231 o The planning of the LEOP sequence and its supporting operations for a mission such as is a time consuming and complex activity. A conservative LEOP plan with time redundancy would be able to absorb minor operational delays caused by transient anomalies either on-board or on-ground. This does not guarantee a timely LEOP but it does prevent a small problem from becoming a critical anomaly that requires replanning (in a period of high stress), causes lengthy delays and essentially means a longer LEOP. o As a complementary measure, a healthy margin in the propulsion supply of the Propulsion Module and in the power subsystem would help to cope with the bigger anomalies that do cause unavoidable delays in the apogee raising sequence, the consequences of which would be an orbit degradation due to air drag (and other perturbations) requiring additional ARBs and a longer stay in the eclipse environment of LEO. Constraints on the satellite sub-systems that limit when/where/how a burn can be made shall be kept to a minimum to allow as much freedom as possible for the Flight Dynamics team when planning the LEOP sequence (or re-planning it, if required). If possible, the perigee should be raised to its operational altitude early on in the orbit raising sequence so as to improve the probability of performing the ARBs in visibility Option of beacon monitoring The use of a Beacon Monitoring system has been considered for use during LEOP: With a Beacon Monitor in use during LEOP the ground operators could have a much longer duration of visibility on the health of the satellite. Small radio antennas around the globe would be sufficient for the reception of the beacon and would mean, for example, that the ARBs would not be performed completely in the blind (unless the orbit perigee is over the middle of the Pacific where there would be no fixed antenna to receive the beacon). Without in-flight calibration (no time for it during LEOP) the beacon could only be relied upon for simple status information but this would be enough for the operators to be able to infer whether an ARB had gone ahead or not, or whether there is an anomaly that threatens the next ARB etc. There would be a delta cost associated with the development of the Beacon Monitor and the ground network would need to be established. The limited benefits of a Beacon Monitor for are significantly outweighed by its development costs and will not be assumed for the operations concept Two satellites solution (Baseline) The ESOC ground segment consists of: The Ground Stations and the Communications Network The Mission Control Centre (infrastructure and computer hardware) The Flight Control System (data processing and Flight Dynamics Software) Infrastructure (Mission Control System, Simulator, etc) The ground segment shall provide: S-band payload data acquisition during the coronagraph operations phase

184 Page 182 of 231 A satellite monitoring and control chain, which includes: o An S-band Housekeeping TM acquisition and processing functional chain o An S-band TC generation and uplink functional chain o Offline performance analysis functions. An orbit and attitude monitoring and control/assessment functional chain An overall Mission Planning function An OBSM facility Data archiving A Flight Control Team (FCT) will be established to execute the operations of the following subsystems: AOCS, Propulsion Module, FF elements (sensors, actuators, FF Executive), Power, Thermal, Communications, Data Handling, and the Coronagraph payload. Also, for the support of the shift work at times of critical operations (LEOP and FF tests), a B Team will be established at LEOP-9 months. B Team positions will be filled by the Simulations Officer and engineers from other Observatory/Survey missions. Please refer to the following subsection, Mission Phases, for the duration of each phase and level of operational support to be expected for each phase. The following text is a discussion on the support to be provided during routine coronagraph science operations Routine operations support The operational orbit has been optimised for visibility from Perth; the Mission Operations Centre (MOC) is located on the other side of the world at ESOC; for someone on the ground the orbit appears to precess around the sky in phase with the passing year (that is, it is fixed in the inertial reference frame); and any ground station pass that might be possible around the 2 hours of perigee would be too short to be of any use. This means that all-year-round daytime-only operations for the routine phase are only possible if, at a certain time of the year, the working day is stretched to either start at 7am or end at 7pm in the extreme. This is when the perigee of the orbit is during the Perth nigh time and the 4h coverage gap happens during the MOC daytime. At this time, starting the working day a little bit earlier or ending it a little bit later allows the daily telecommunications period (DTCP) to be split into two parts: taken from the start of the MOC day until the Perth Loss Of Signal (LOS), and then starting again 4 hours later from the Perth Acquisition Of Signal (AOS) until the end of the MOC day. For the rest of the year the DTCP window can remain whole within the MOC daytime. For Perth scheduling, the DTCP window will progressively shift throughout the year. In terms of the orbit, the DTCP window will progressively rotate around it such that after the 18 months of coronagraph science observations the window will have gone round approximately 540º. This all relies on the correctness of the assumption that a high-rate data downlink (250kbps) is possible at any time during the 20.4hrs visibility period from Perth, regardless of the activities going on (such as FF manoeuvres or coronagraph science observations). Regardless of this, the absolute baseline for the routine phase is to provide on-console SPACON support for a 6h DTCP and full engineering support during office hours. If the DTCP is out of office hours, such as over the weekend or during the week in the event that the DTCP can only be at a particular point in the orbit, then engineering support will be via on-call support.

185 Page 183 of 231 Flight Dynamics will not provide orbit control as in the classical sense (although they will still make use of the range and Doppler measurements taken during the DTCP). There is not enough time to perform range and Doppler measurements immediately after perigee, do the orbit determination, calculate the FF manoeuvre commands and uplink them in time to begin scientific observations. Instead, on-board orbit determination using GNSS measurements taken around perigee will have to be relied upon along with the capability of the on-board FF Executive to control the formation from its construction to its de-construction. Flight Dynamics will be responsible for the: Formation Design and its integral management of propellant consumption Relative Navigation and Formation Control analysis and assessment (in co-operation with the Flight Control Team) Performance analysis of the on-board orbit determination from the GNSS products. Entirely new software will have to be created for these tasks which should be able to serve as the basis for software for other FF missions to come. In addition to these FD software developments new simulator FF models and calibration support software for the FF Executive (linked with the FD s/w) to be the first generation of a class of software products that can be inherited by later missions can be expected Mission phases The following table provides additional information about the operations activities according to mission phase as well as the level of ground station support that will be required. Mission Phase Duration Operations Support Operations Definition Phase Operations Implementation Phase Concept and Requirements Definition Phase Design and Development Phase Up to L 4 years L 4 years to Launch L 4 years to L 3 years L 3 years to L 6 months Operations concept definition, special developments (e.g. Radioscience developments), mission analysis See subphases below: Requirements Definition Phase, Design and Development Phase, and Test and Validation Phase The operations concept and the orbital strategy are detailed. Detailed project documentation (e.g. user manual and data bases) is received. All the operational tools are developed, the Flight Operations Plan is generated, the teams as well as the network and stations are set up. Listen-In tests and SVT 0, 1 and 2 may be performed. Ground Station(s) and On Board Antennas

186 Page 184 of 231 Mission Phase Duration Operations Support Test and Validation Phase LEOP Commissioning and Verification L 6 months to Launch L 8hrs to L + 12 days L + 12 days to L + 42 days Final mission simulations and training, final Flight Operations Plan and SVT 3. (remark: number of SVTs is an assumption for costing purposes) Real-time shift support (satellite operations, flight dynamics, ground segment maintenance and support). At least three S-band ground stations used, giving as much daily coverage as possible. No payload operations support Verification of the basic satellite configuration. Preparation of the satellite for, and execution of, the orbit raising manoeuvres. Real-time support for satellite and payload operations (presence of the PI Team with decision authority required at ESOC). Off-line flight dynamics support. On-call ground segment maintenance support Off-line support for manual operations re-planning, data analysis, troubleshooting. One station only used throughout the phase. PISA for PI Team to be provided. Optimised Commissioning Plan for the classical subsystems only, followed by the first set of FF tests (incl. FF subsystems commissioning). Instrument commissioning. Satellite system tests including FDIR and redundancy tests. Pointing, Calibration and Performance campaign for the STR and redundant sensors. Ground Station(s) and On Board Antennas ESTRACK 15m S-band as required, use of LGA on board Perth 15m S- band use of LGA on board

187 Page 185 of 231 Mission Phase Duration Operations Support Routine (Coronagraph Science Observation) Phase L + 42 days to end of mission (L + 20 months) FF test days 12 in total (2 sets of 6 noncontiguous days). 1 st set performed during Comm/PV. Remaining days spread throughout the Routine Phase. Off-line operations support (satellite operations, flight dynamics, ground segment maintenance) Daily contact of 6 hours, automatic pass operations. Payload operations ruled by routine mission planning. Priority setting according to Master Science Plan. One station only used. Range and Doppler measurements taken. On-line operations support (satellite operations, flight dynamics, ground segment maintenance) Daily contact of 20 hours, manual operations. Range and Doppler measurements taken. Table 19-1: Mission phases overview Ground Station(s) and On Board Antennas Perth 15m S- band LGA Perth 15m S- band NNO 35m S- band backup LGA Ground stations and communications network The ground stations network to be used for during LEOP will be composed by the 15m antennas in Kourou, Villafranca and Perth plus other S-band antennas selected to improve the ground visibility. For the routine coronagraph science phase the 15m antenna in Perth is the baseline. For the FF test days the 35m antenna at New Norcia is the baseline backup depending on the use of this station by other missions. The test days shall be chosen so as not to conflict with other missions. The Ground Facilities Control Centre monitors and has the ability to remotely control all the ESTRACK ground stations, using information provided by Flight Dynamics and the ESTRACK scheduling facility. They will also be responsible for the TM/TC links to and from the ground stations and any data retrieval of stored coronagraph science data from the TPS or the auxiliary data used by FDS. A station computer monitors and controls (locally, automatically or remotely from the operations control centre) all equipment on the station. A Front-End controller unit controls the antenna subsystem.

188 Page 186 of 231 All computing systems in the Mission Operations Centre (MOC) will be connected by a Local Area Network (LAN) to allow transfer of data at sufficient speed and to allow joint access to data files. The Operations Computer Systems (the machines for the S2K server and clients, mission planning, flight dynamics, network interface and ground segment control) will be connected by the OPSLAN. The OPSLAN and the ground stations are connected via the OPSNET, a closed Wide Area Network for data (telecommand, telemetry, tracking data, station monitoring and control data) and voice The mission control centre The P3 mission will be operated from ESOC and it will be controlled from the MOC, which consists of the Main Control Room (MCR) augmented by the Flight Dynamics Room (FDR), Dedicated Control Rooms (DCRs) and Project Support Rooms (PSRs). The MCR will be used for mission control during LEOP and possibly the Commissioning Phase in case of serious anomaly. During cruise and the asteroid operations phase the mission control will be conducted from a Dedicated Control Room shared with other Interplanetary missions, such as BepiColombo and Rosetta. The control centre is equipped with workstations giving access to the different computer systems used for different tasks of operational data processing. The control centre will be staffed by the Flight Control Team (FCT), with support from operations engineering staff, experts in spacecraft control, flight dynamics and network control, available on a part time basis for the full mission duration. Space and equipment for scientists, project and industry experts and public relations will be provided close to the MOC as required, during the critical phases of the mission. The Coronagraph Science Operations Centre will be located at the European Space Astronomy Centre at Villafranca, Spain Computer facilities The computer configuration used in the MOC for the mission will be derived from existing structures. The computer system basically consist of A computer system used for the Flight Operations Plan generation in a form directly usable by the mission dedicated computer A mission dedicated computer system (including workstations hosting SCOS2000 (S2K)) used for real time telemetry processing and for command preparation and telemetry and command log archiving, and also for non real time mission planning and mission evaluation Workstations hosting the flight dynamics system The simulation computer, providing an image of the satellite system during the later phases of ground segment verification, for staff training and during operations Computers for the data dissemination infrastructure elements The network interface system The ground segment control system. All computer systems in the control centre will be redundant with common access to data storage facilities and peripherals. In particular, the prime and redundant S2K servers will be physically

189 Page 187 of 231 split. Preferably workstations of a similar type will be used for all related computing, to maximise flexibility and to minimise maintenance costs. The work stations hosting the S2K clients will be located in the different control rooms as necessary, and the clients configured with the different levels of user privileges as required The flight control software system The Flight Control System will be based on infrastructure development (S2K), using a distributed architecture for all satellite monitoring and control activities. The Flight Control System includes the following facilities: Telemetry reception facilities for acquisition, quality checking, filing and distribution Telemetry analysis facilities for status/limit checking, trend evaluation Telecommand processing facilities for the generation of commands for control, master schedule updates, and on-board software maintenance. The facilities will provide also uplink and verification capabilities Monitoring of instrument housekeeping telemetry for certain parameters that affect satellite safety and command acceptance and execution verification Separation and forwarding of payload telemetry to Science Data Processing Centres Checking, reformatting, scheduling command request for payload. Within the S2K system, mission specific software will be developed wherever necessary Mission Operations Concept Overview The mission operations will comprise: Satellite Operations, consisting of mission planning (where necessary), satellite monitoring and control and all orbit and attitude determination and assessment/control of the Coronagraph and Occulter satellite. Mission Operations proper will commence at separation of the satellite stack from the launcher and will continue until the end of the mission, when ground contact to the satellite will be aborted. Mission Operations will comprise the following tasks: Mission Planning: long term planning and short term planning (24 hours to one week time frame) Satellite status monitoring Satellite control, based on monitoring and following the FOP and the short-term plan Orbit determination and control using tracking data and defining manoeuvres Attitude determination and assessment/control based on the processed attitude sensors data in the satellite telemetry and by commanded updates of control parameters in the onboard attitude control system On-board software maintenance Maintenance of ESA ground facilities and network.

190 Page 188 of Mission planning, satellite monitoring and control The Operations support activities for will be conducted according to the assumptions in section 19.2 and can be summarized as follows: All operations will be conducted by ESOC according to procedures contained in the FOP Nominal Satellite control during the routine mission phase will be off-line. The contacts between the Mission Control Centre and the Satellite, except for collecting payload and housekeeping telemetry, will therefore primarily be used for pre-programming of autonomous operation functions on the satellite, and for data collection for off-line status assessment. Anomalies will be normally detected with some delay All operations will be conducted by the uplinking of a master schedule of commands for later execution on the satellites. The master schedule will be prepared by a dedicated Mission Planning System. The Mission Planning System will be also used for planning payload operations During the LEOP phase TT&C S-band operations will be conducted from the ESOC MCR During the routine phase, operations will be conducted from an ESOC DCR Orbit and attitude control The Flight Dynamics support will consist of: Orbit determination of the satellites during the LEOP and routine phases using two-way Ranging and two-way coherent Doppler data (TBC) Manoeuvre optimisation: the manoeuvres performed for apogee raising will be optimised to minimise propellant consumption and taking into account all operational conditions Formation Design and its integral management of propellant consumption Relative Navigation and Formation Control analysis and assessment (in co-operation with the Flight Control Team) Performance analysis of the on-board orbit determination from the GNSS products Attitude Control System Monitoring: monitoring and verification of the on-board functions such as star tracker window and sensitivity setting Manoeuvre command generation: preparation of command sequences for input to the master schedule updates related to all orbit and attitude manoeuvres Manoeuvre monitoring Calibration of thrusters and sensors.

191 Page 189 of PROGRAMMATICS/AIV 20.1 Requirements and Design Drivers The following requirements and design drivers have been used: 2 satellites (baseline): Coronagraph Occulter 3 rd satellite (optional): identical to Occulter Launch date: Mid 2010 Operations phase: 6 month 2 years Launch: VEGA (Eurorockot/DNEPR optional) 20.2 Assumptions and Trade-Offs All satellites are structurally similar, that is, qualification is achieved with one structural-thermal model (STM). The Coronagraph includes two small deployable solar array panels, while the other satellite has only fixed wing solar arrays. The Coronagraph also includes a deployable cover for the instrument. The alignment requirements for all satellites are rather demanding. The satellites shall be built, as far as possible, in parallel. This has an impact on the required MGSE and EGSE. A number of items will be needed in multiple (for example integration stands) and only a limited number of items will be retrieved from other projects. Environmental acceptance testing will be performed, as far as possible, for all satellites at the same time. The project development is assumed to be performed by an industrial team whereby the Prime Contractor is responsible for the: Overall mission analysis Overall design development and procurement of the satellites Detailed satellite design at system and subsystem level Direct procurement of the satellite units, equipment and major assemblies (hardware and software) Overall satellite Assembly, Integration and Verification (AIV) activities Definition and control of the technical and operational interfaces of the Instruments.

192 Page 190 of Two Satellites Solution (Baseline) Master schedule Figure 20-1: Master schedule Model philosophy Considering the moderate development risk identified for most aspects of the satellite design, a Protoflight approach has been selected at satellite level, based on a four models philosophy: Structural Thermal Model (STM): will ensure the mechanical and thermal qualification of the satellite main structural design. Most of the unit assemblies will be represented by structural dummies Avionics Test Bench model (ATB): will ensure verification of the overall electrical, functional and software interfaces. Breadboard units (BBs) will be used most of the time, exceptionally Interface Simulators could be used for the Payload Units. Elegant BB units (EM like with commercial components) or modified EM could be used if cost-effective, for example in case of recurring units with EM available, or off-the-shelf equipment Protoflight Model (PFM): The Coronagraph satellite will follow the Protoflight concept at system level, but will reuse the STM. Delta qualification at system level might be necessary; in that case selected test will be performed at qualification test levels, but with acceptance duration Flight Model (FM 2): Built to full flight standard will be subjected to acceptance test levels for acceptance duration.

193 Page 191 of Model and test matrix Test Description STM ATB PFM FM 2 Mechanical Interface R, T - R, T Mass Property A, T T T Electrical Performance - T T T Functional Test - T T T Deployment Test A, T A, T - Telecommunication Link - T T Strength Load T Q - - Shock T Q - - Sine Vibration T Q T A T A Modal Survey Acoustic Noise T Q T A T A Outgassing A, I - - Thermal Balance A, T Q T A - Thermal Vacuum T Q T A T A Grounding and Bonding - R, T V R, T V EMC Test - T V T V Magnetic Test Table Abbreviations: T A : Test at acceptance level T D : Development tests T Q : Test at qualification level T V : Verification by Measurements A: Analysis E: Equipment level I: Inspection R: Review of design Table 20-1: Model and test matrix AIV plan Taking into account the given model philosophy and the expected development time of the Instruments, an overall AIV planning is outlined below. The validity of the planning is based on the following key assumptions: The instruments mechanical and thermal design is qualified by the suppliers, including verification of inner optical alignment stability. Final qualification will be achieved on the satellite PFM model Instruments STM models will be delivered at a build standard compatible with the satellite STM programme. Mechanical alignment only will be verified at satellite level Instruments BBs will be delivered at a built standard which, as a minimum has to be representative of the electrical and functional interfaces. No calibration activities are envisaged on the ATB

194 Page 192 of 231 The Instruments PFM and FM will be installed on the satellites by the prime contractor. Functional and interface tests will be performed to certify their proper on board accommodation. Calibration activities will be performed on all flight satellites. Figure 20-2: AIV Plan Summary and critical issues The integration of the satellites imposes some critical assumptions. The first is that to perform a timely delivery of the flight units, their test activities have to be performed partially in parallel, as shown in the schedule. It is possible to minimise the effect of parallel work applying the following criteria: The integration of the flight models is performed in sequence. Therefore, one integration team and one set of integration tools are necessary. The environmental tests are performed in the same facilities, and in sequence on the two satellites. This way, one set only of test adapters is needed. Duplication of the MGSE is envisaged to handle the two flight satellites. Two containers and two integration dollies are needed. Some savings are expected on the other MGSE items, where the discontinuous utilisation allows for shared use.

195 Page 193 of 231 The same principles of satellite integration and testing are applied to the structure and propulsion subsystem AIV. Assembly and test of the structure models are done in sequence, to avoid duplication of tools and teams. The same applies to propulsion modules. Another critical aspect is the timely release of the SW versions. The on-board SW first version V 1.0 must be ready for the start of the ATB activities. The final version, implementing the results of system testing on ATB, must be loaded on the PFM units before they are delivered to the system for integration. This final release V 2.0 shall be also tested on the ATB. This late test will give the final confirmation of adequacy of SW implementation of the system functions. The desired launch date in mid 2010 can be achieved with the above assumptions, even with some margin. A launch date in March 2010 is considered feasible Three Satellites Option This investigation was focussed initially on a three satellites solution. However, for various reasons this solution was abandoned and the two satellite solution became baseline. The results of the three satellites investigation are presented for information Master schedule (3 satellites option) Figure 20-3: Master schedule Model philosophy The model philosophy for the three satellites option would be the same as for the two satellites baseline. FM 3 is an identical rebuild of the FM Model and test matrix The test matrix for the third satellite (FM 3) would be the same as for FM AIV plan The AIV plan for the third satellite is basically a copy of the planning of the Occulter satellite. By using this strategy, the third satellite will be automatically on the critical path instead of the second satellites as shown below. The key assumptions are the same as for the two satellites baseline.

196 Page 194 of 231 Figure 20-4: AIV Plan (3 satellites option) Summary and critical issues The integration of the three satellites imposes the same critical assumptions as for two satellites. The same criteria will have to be applied as described in chapter This brings the end of the tests on the PFM at the beginning of the tests on FM 3. Thus the FM 2 is tested in parallel with the PFM first and then only for a limited time also the FM 3. The consequence is that also for the three satellites option only two sets of EGSE are needed. The desired launch date mid 2010 could be achieved with the above assumptions.

197 Page 195 of COST 21.1 Requirements The responsibility of the CDF cost chair is to estimate the total industrial cost. However, considering the circumstances of the mission, in accordance with the team leader and the customer, all the mission costs have been considered, with exception of the Ground Segment & Operations. For completeness, the Ground Segment and Operations estimate from ESOC has been included in the summary tables of this chapter. presents a very challenging case from the cost estimate point of view as no specific heritage at system level exists within ESA missions. According to the ESA Cost Engineering Office Chart of Services, the estimate performed is at class 4 level (RD[78]) Assumptions Organization The PROBA missions are managed by a project team within the Systems Department of ESA (TEC-S). is prospected to be an order of magnitude more expensive and much more complex than the previous PROBA missions. The ESA project team in TEC-S has been retained as main assumption for the project environment. In order to describe the global organization, the following elements of the Life Cycle Cost (LCC) are identified: System, all tasks that go to industry Launch and Phase E, including transport, commissioning, industry support and operations ESA internal cost The elements are assumed to be hierarchically organised as in Figure ESA System Launch and Phase E ESA Internal Cost System Prime Activities Coronagraph Science instrument Launch (VEGA) LEOP Ground Segment and Operations Technology Developments and GNC Algorithms Coronagraph satellite Occulter satellite Optical Metrology Figure 21-1: assumed project organization

198 Page 196 of 231 The System Element includes also the coronagraph instrument, which is assumed as Furnished Equipment, so not included in the cost estimate, but the satellite AIT includes the costs of its integration Cost Breakdown Structure (CBS) The CBS at its lowest level is shown in Table 21-1 together with the assumed inputs for the independent cost estimate. The Technology Readiness Levels (TRLs) are defined in RD[80]. In particular, the following assumptions are highlighted: The Model Philosophy at system level is in line with the work presented in the Programmatics chapter. The Propulsion Module and the launcher adapter are Off-The-Shelf from the ESA LPF Project, which, at the time of writing this report, is at the beginning of Phase C/D. As a consequence, it is assumed at TRL 4 today and TRL 6 by the beginning of phase C/D. Minor or no modifications are assumed with respect to LPF. Optical Metrology. o In Table 21-1 the Optical Metrology Instruments are accounted separately from the Occulter satellite, however, the AIT is included in that satellite AIT costs. o In order to reduce the cost, it is assumed to develop only mass dummies for the satellite STM, 1 Engineering Model and 1 (or 2 in the three satellites option) Engineering Qualification Models to be refurbished and flown in the satellite Flight Models. The models, according to RD[79], will be used for functional qualification (EM) and functional performance qualification (EQM); they may be equipped with commercial parts, replacing the most expensive hi-reliability parts. o The TRL of these instruments is obtained from the team specialist. o Very low level of details drives a low Level of Confidence on the cost estimates. Inter-Satellite Link (ISL) is assumed to be developed from a Commercial Off-The-Shelf equipment, properly modified for the space environment. Expensive space qualified hireliability parts are not foreseen. This equipment is considered only for the optional concept with three satellites. The proposed RF metrology equipment is from a pre-development of Alcatel that successfully built a breadboard under ESA funding. GSTP-4 covers additional funding to go to EBB and CNES will fund an EQM. The equipment will be used in the PRISMA mission to fly in However, some modifications for will be necessary because PRISMA does not include requirements for collision avoidance and 4-π visibility. In the frame of, it is assumed to build 2 new EQMs or QMs, further test them and fly them refurbished Project environment The industrial cost is estimated with a large contractor (EADS Astrium or Alcatel Alenia) as System Prime and two Medium/Small contractors (for example SSTL, OHB, Carlo Gavazzi Space) in charge of building the platforms. That is retained as most likely hypothesis,

199 Page 197 of 231 considering the funding situation. Also, the possibility to have a single company acting as a Prime and procuring equipments to assembly in two recurring platforms, is taken into account as opportunity in the Cost-Risk Analysis.

200 Page 198 of 231 Mission Cost Breakdown Unit Mass Performance TRL Design Number H/W Matrix kg Name [Unit] Value Today Ph. C/D start Status of Units MD STM EM EQM QM PFM FM Sp ESA TEC-SYC Estimate # x Technology developments GNC Algorithms and SW 2-1 RF Metrology 4-1 InterSatellite Link No 3-1 Optical Metrology 4-1 FF Validation Bench 2-1 Micro-propulsion 4-1 Total Technology developments Platform A - Coronograph Phase B-C/D HW-SW AOCS Star Trackers 0.2 FOV [deg] Coarse Sun Sensors 0.2 FOV [deg] Gyro 0.2 Accuracy GPS receiver Wheels 1.2 Momentum [Nms] RCS - Cold Gas Propulsion Thrusters 0.1 Thrust [m N] 40 / Tanks 12.7 Volume [dm3] Valves and piping 6.0 Propulsion type CG k AIV Electrical Power S/S Solar panels 1.8 Total Area [m2] k SADA Max Power transf. [W] Battery 5.8 Capacity [Wh] PCDU 6.6 Max Power Conv. [W] k TTC and RF metrology S-Band system Band S k Inter Satellite Link No RF metrology system k Data Management System CDMS 6.5 No. boards k On-board Software Structures, Harness and Thermal Structures 21.2 Complexity Simple Harness S/C Volume k Thermal Control 4.0 Controlled Volume [dm3] k Phase A-B-C/D system activities Phase A-B-C/D Project Office Duration [years] 4 Coronograph AIT&V RACE Complexity 2 Coronograph GSE RACE Complexity 3 Total Coronograph S/C Platform B - Occulter Phase B-C/D HW-SW AOCS Star Trackers 0.2 FOV [deg] Coarse Sun Sensors 0.2 FOV [deg] Gyro 0.2 Accuracy GPS receiver Wheels 1.2 Momentum [Nms] RCS - Cold Gas Propulsion Thrusters 0.1 Thrust [m N] 40 / Tanks 12.7 Volume [dm3] Valves and piping 6.0 Propulsion type CG k AIV Electrical Power S/S Solar panels 1.9 Total Area [m2] k Battery 5.8 Capacity [Wh] PCDU 6.3 Max Power Conv. [W] k TTC and RF metrology S-Band system k Inter Satellite Link No RF metrology system k Data Management System CDMS 6.5 No. boards k On-board Software kloc Structures, Harness and Thermal Structures 21.2 Complexity Occulter Disc Harness k Thermal Control 4.0 Controlled Volume [dm3] k Phase A-B-C/D system activities Phase A-B-C/D Project Office Duration [years] 4 Occulter + Optical Metrology AIT&V Complexity 3 Occulter GSE 3 Total Occulter S/C Optical Metrology instruments Included Coarse Lateral Yes Project Office FTE [ Manyears] 3 AIT+GSE Transm itter Receiver Electronics Fine Lateral Yes Project Office FTE [ Manyears] 4 AIT+GSE Transm itter Receiver Electronics Absolute Longitudinal (DWI) Yes Project Office FTE [ Manyears] 6 AIT+GSE Lasers system Optical bench Electronics Fringe Tracking Sensor (PDS) Yes DW I add-on Project Office FTE [ Manyears] 4 AIT+GSE Laser Optical bench Electronics (add-on) Total Optical Metrology System Prime Costs Propulsion Module and Launcher Adapter Coronograph System Simulator and Verification 2-1 Phase A-B-C/D Project Office Duration [years] 5.5 Total System Prime Costs Table 21-1: System CBS

201 Page 199 of Two Satellites Solution (Baseline) Baseline cost estimate Mission Cost Breakdown Cost Estimates ESA TEC-SYC Estimate % absolute % relative Proba3 CDF 2 S/C concept Total ESA Internal Cost 6.0% Phase E Launch (VEGA) 60.0% LEOP (incl. Ind. Supp. + Transport, Integration, Commissioning) 11.4% Ground Segment & Operations 28.6% Total Launch and Phase E 23.5% System Total Technology developments 15.1% Total Coronograph S/C 17.6% Total Occulter S/C 18.8% Total Optical M etrology 15.5% Total System Prime Costs 18.9% Design and Cost Maturity Margin 17% 14.2% Total System Cost 70.5% Total Cost Core Estimate 100.0% Table 21-2: Total Cost Estimate Summary The cost of Phase E largely depends on the launch cost and the Ground Segment. It should be noted that there might be an opportunity to launch with one of the first launches of VEGA, which may be covered under the VERTA programme, and reducing dramatically the charge on the project. That is included into the Cost-Risk (see later); the figure included is a provision based on currently available information. The cost of Ground Segment and Operations included is also a provision, as ESOC is normally in charge of the estimate. The ESOC estimate certainly overrules this provision. The Design Maturity Margin, as explained later, is a cost that is expected to occur, but cannot currently be included in the estimate due to the low design maturity. The project team assumed is of a reasonably fair size, but the actual team depends largely on the Project Manager in charge and the key personnel, that on individual level may have specific or generic skills that may drive a particular team composition. The heavy industrial consortium assumed unavoidably impact the cost with duplicated functions. It is acknowledged that some of these functions are oversized, but it is believed that with the current set-up it cannot be avoided. In fact, the estimated effort is to be considered optimistic and stringent ESA requirements may lead to significantly larger team sizes.

202 Page 200 of 231 Higher cost is expected, but in order to keep the fair estimate to certain limited figures (also considering the conservative approach on the organization) it was decided to put the possible higher cost as a Cost-Risk item. The Software development is an important part of the development. The design maturity does not allow for a detailed characterization of the work, however, with the hypothesis of using existing space technologies, the mission software can be assumed as follows: FF Control Loop Algorithms o Physical characterization o Service Layer Software On-board Software o GNC - Applications Software - Basic Software (hardware drivers) o DMS Ground Support and Validation o ISVV o Mission Simulator The peculiarity of is assumed to be concentrated in the first bullet. The physical characterization is the model of the FF behaviour under all possible conditions and inputs in closed control loop; the Service Layer SW is the translation of the equations into a code that generates commands to control the system, based on generic hardware I/O. Some estimate details are included in the following figures. Platform Cost Breakdown GSE AIT Project Office Structures, Harness and Thermal Data Management System TTC and RF metrology Electrical Power S/S RCS - Cold Gas Propulsion AOCS Platform A - Coronograph Platform B - Occulter Figure 21-2: Platform Core Cost Estimate Breakdown

203 Page 201 of 231 Optical Metrology Fringe Tracking Sensor (PDS) 34% Coarse Lateral 14% Fine Lateral 16% Absolute Longitudinal (DWI) 36% Figure 21-3: Optical Metrology Cost Breakdown Cost-risk analysis Experience and lessons learned show that, statistically, the probability distribution of cost estimates is skewed to the left, making the most likely estimate more distant from the possible maximum value, than from the possible minimum value. Consequently, the algebraic summation of most likely estimates is not equal to the most likely total. Furthermore, the total cost at a given Level of Confidence has to be calculated with simulations that take into account correlations among the cost elements. ESA TEC-SYC has developed a procedure to estimate the most likely total and to give a result at a specified Level of Confidence, based on a Monte Carlo simulation and spreads of possible costs based on four categories (RD[81]). The procedure is summarized in the following table.

204 Page 202 of 231 Table 21-3: Cost-Risk Contributors 2 The margins and the total cost at the given Level of Confidence are calculated using a Monte Carlo simulation with the It should be noted that the Cost-Risk Analysis, by considering the main risks associated to the project (both technical and programmatic) and their statistical impact on the total cost, provides a systematic approach to the identification of the contingency margins to be applied to the programme budget. Project owned Cost-Risk and Project Environment Cost-Risk thus represent different contingency allocations that should be budgeted in order to be able to stay (with a 70% confidence level) within the total Budget at Completion. Given the different nature of the costrisks, it is recommended to identify different contingency owners (for example Project level, Programme level). On this basis, it is important to realise that the figures referring to the Cost-Risk Volumes and related Cost-Risk Contingencies are to be treated separately from the core estimate and the expected cost of the industrial proposal for the implementation phase. For the same reason, it is also of paramount importance that such contingency figures are not disclosed outside the ESA executive. The assumed cost-risk characterization is shown in Table The calculated EPE margin is not intended to be exhaustive and may be enriched by additional items specific to the mission.

205 Page 203 of 231 Mission Cost Breakdown Cost-Risk Inputs 95% 90% 80% Cost Model Design 110% 130% 150% ESA TEC-SYC Estimate Accuracy Maturity Project Ow ned Events Score Proba3 CDF 2 S/C concept + / - + L M H Technology developments GNC Algorithms and SW H RF M etrology L InterSatellite Link No M Optical Metrology M FF Validation Bench L Micro-propulsion L Total Technology developments Platform A - Coronograph Phase B-C/D HW-SW AOCS L RCS - Cold Gas Propulsion L Electrical Power S/S L TTC and RF metrology L Data Management System and Appl. SW M Structures, Harness and Thermal L Phase A-B-C/D system activities Phase A-B-C/D Project Office M Coronograph AIT&V L Coronograph GSE L Total Coronograph S/C Platform B - Occulter Phase B-C/D HW-SW AOCS L RCS - Cold Gas Propulsion L Electrical Power S/S L TTC and RF metrology L Data Management System and Appl. SW M Structures, Harness and Thermal L Phase A-B-C/D system activities Phase A-B-C/D Project Office M Occulter + Optical Metrology AIT&V M Occulter GSE L Total Occulter S/C Optical Metrology instruments Included Coarse Lateral Yes M Fine Lateral Yes M Absolute Longitudinal (DWI) Yes H Fringe Tracking Sensor (PDS) Yes H Total Optical Metrology System Prime Costs Propulsion Module and Launcher Adapter M Coronograph System Simulator and ISVV M Phase A-B-C/D Project Office M Total System Prime Costs Table 21-4: Cost-Risk characterization The analysis shows that significant budget reductions are possible through opportune decisions. In particular, in the frame of the EPE, the choice to go with a single Prime Contractor can bring important cost savings. Furthermore, there are two major opportunities for reducing the cost: Enhance competition In case of two different platform integrators, establishing a central procurement entity. The results of the analysis are shown in the following figure:

206 Page 204 of 231 Figure 21-4: Cost-Risk Margins Contributors The choice of the opportune Cost-Risk Margin is left to the Project Manager to decide upon. The Cost-Risk analysis is a process evolving with the design maturity and the profiles shown in Figure 21-5 can be expected. The Risks decrease with the maturity of the project and some effects, and some actions translate in additional costs to be added into the core estimate. Figure 21-5: Cost-Risk Evolution

207 Page 205 of 231 It should be clarified that while CMA, POE and EPE costs may or may-not occur, the figure estimated for the Design Maturity (DM) is expected to translate fully in project costs, and for that reason it is part of the Core estimate Three Satellites Option The CDF study has evaluated the possibility to design a mission with three satellites, with enhanced FF demonstration. In that concept the following delta cost apply, assumed that: The third satellite is a recurring model of the Occulter (including the metrology instruments). All satellites include an InterSatellite Link (ISL). The Coronagraph satellite remains unchanged, except for the introduction of an ISL. The Control loop algorithms are significantly more complex, as well as the system simulator and validation bench. The master schedule will not change significantly (~3 months). The launcher and upper stage (LPF PM) remain the same. 3 rd Satellite Delta Cost Platform 5.7% System Costs 2.1% ISL 0.6% Optical Metrology 6.5% GNC control loop and SW 2.9% System Validation 0.7% SubTotal 18.6% Ground Segment and Ops 3.4% Delta schedule 1.0% Total delta cost 23.0% Table 21-5: Delta Cost of third satellite

208 Page 206 of 231 This Page Intentionally Blank

209 Page 207 of CONCLUSIONS This chapter highlights the overall conclusions and critical aspects resulting from the study performed by the CDF team, in addition to the discipline specific conclusions which have been reported in the previous chapters. The following two sections provide the compliance of the study results, mission and system design with the study and mission objectives Compliance with Study Objectives Table 22-1 provides the compliance for each of the study objectives, provided at the beginning of the study (refer to section 2.2). Study Objective Review the Formation Flying (FF) technologies requiring in orbit demonstration and the related requirements Assess the technical feasibility of a FF demonstration mission fulfilling the payload requirements within the given constraints Identify the best compromise between the number of, and the level of, demonstrated FF technologies and mission complexity, cost, and the interest for future ESA missions Identify the derived mission and system requirements Present a preliminary system baseline design, including trade-offs, analysis and the identification of critical aspects and areas Present a preliminary formation flying mission implementation plan and cost envelope identification For the de-scoped FF technologies, provide an indication of the rationale and of their impact on system design Table 22-1 Study Objectives Compliance Compliance A review of the customer provided information on FF has been performed. Wherever relevant this is reported in the discipline chapters. Within the given constraints, the technical feasibility of a demonstration mission for FF has been shown. Sub-system critical issues have been identified and reported. Some system and mission critical issues have been identified and reported. The study results provide information on the compromise between the requested parameters The system requirements derived during the study are reported in the Systems chapter A system baseline has been defined including options and trade-offs at system and subsystem level. Critical design issues have been highlighted. The Programmatics aspect and cost estimate of the baseline mission have been provided. The industrial and Operation cost are reported in a separate document, CDF-42(B). In the Systems as well as the different subsystem chapters, a rationale has been given to which technologies were used, and which were de-scoped.

210 Page 208 of Compliance with Mission Objectives Table 22-2 provides the compliance for each of the mission objectives, provided at the beginning of the study (refer to section 3.2). Mission Objective Compliance Flight qualification of FF hardware, which are currently foreseen to be: RF metrology Chapter 6 reports on the technology and the implementation used in the preliminary baseline design Optical metrology Chapter 7 reports on the technology and the implementation used in the preliminary baseline design High accuracy propulsion system Chapter 11 reports on the technology and the implementation used in the preliminary baseline design Demonstration of FF manoeuvres for Chapters 5 and 9 discuss the demonstration future missions of FF manoeuvres, also in relation to future Exploitation of the FF for Scientific return Table 22-2 Mission Objectives Compliance missions Chapter 8 presents the coronagraph instrument used in the preliminary baseline. The resources necessary for the operation of the instrument have been allocated in the various subsystems, providing observation duration of approximately 18 months Compliance with Mission and Programmatics Constraints The Mission and Programmatics constraints were provided to the study team at the beginning of the study. These constraints reflect the fact that will be a demonstration mission for FF. This implies that the resources are much lower then those of a FF satellite mission for Scientific or applicative purposes. Table 22-3 shows how the Mission and Programmatics have been taken into account during the study at system and subsystem level. The effects of the different constraints are explained in Chapter 5 for Systems and Chapter 20 for Programmatics and AIV.

211 Page 209 of 231 Mission and Programmatic Constraints development and launch shall take place in the near future, not later than 2010, in order to provide a timely feedback for operational ESA missions planned for the next decade. The Operational Phase shall be from 2010 to 2012 The mission shall be designed with a low cost approach, and is assumed to cover: System Analysis Additional development of FF Technologies to achieve flight readiness Ground Verification Procurement of Platforms Assembly, Integration and Test Launch and Operations/Data Analysis GSTP shall provide programmatic framework for the implementation of is planned for initiation in 2006 The implementation shall follow the classical Satellite Project Approach Phase A: Feasibility including enabling technologies up to TRL 4 Phase B: System Definition including technologies predevelopment up to TRL 6 Phase C/D: System Procurement Phase E: System Operation Compliance Constraint fully met Constraint fully met The cost report, CDF-42(B), elaborates the overall ROM cost estimate and the related assumptions and cost risk issues. Constraint assumed during study Constraint taken into account in the study Programmatics (Chapter 20) Constraint taken into account in the study Programmatics (Chapter 20) Table 22-3 Mission & Programmatic Constraints Compliance 22.4 Overall Conclusions and Recommendations A FF demonstration mission including a minimum set of FF manoeuvres, combined with an instrument for Sun Corona observation has been analysed and conceptually designed. The two satellite baseline represents the best compromise between design simplicity, cost, and satisfaction of a minimum set of user requirements. A three satellite option, the use of additional instruments, the use of different propulsion systems have also been traded off in terms of feasibility and fulfilment of requirements and constraints.

212 Page 210 of 231 The design for specific FF key subsystems (GNC in particular) encompasses the use of new and European technologies for demonstration purposes, while the other subsystems have been designed following a low cost and risk approach (using conventional architectures and high TRL components). This coupled with a launch date in 2010 and a maximum two years operation will ensure a timely return of information for exploitation in future ESA operational missions. A preliminary quantification of the demonstration coverage with respect to the future missions is provided in Chapter 9 (GNC). In the proposed baseline, no technical showstopper has been identified. The design and development of all the necessary new technology have already been initiated. Therefore it can be concluded that this FF demonstration mission is feasible. Additional investigation is recommended regarding the launcher and propulsion system in order to further evaluate the alternative options. The embarking of an alternative instrument with more stringent FF requirements might also be considered. Further analysis and refinement of the FF demonstration manoeuvres is also recommended in order to optimise the delta-v usage. Finally, in order to ensure the best return of the demonstration mission, careful consideration should be made to select manoeuvres/technologies that are complementary to those performed/employed in the Prisma and ATV missions. This is particularly true for the two satellites formation.

213 Page 211 of REFERENCES RD[1] LISA Pathfinder CReMA, S2-ESC-RP-5001, MAO WP-450, M Landgraf, Issue 2.0, ESA/ESOC, September 2005, RD[2] Rockot User s Guide, Issue 4.0, November 2004, Eurockot Launch Service GmbH, Bremen, RD[3] Dnepr User s Guide, Issue 2, 2001, Kosmotras, Moscow, RD[4] VEGA User s Manual, Issue 2.0, September 2004, Arianespace, Evry, RD[5] TMTC-Antenna. Design Report Document-No.: PB2-STT-S2023-RPT-100. Issue: 1 Date: STT-SystemTechnik GmbH. Germany. RD[6] Formation Flying RF Subsystem. RF Sub-System Architectural Design. 14 December RD[7] Formation Flying RF Subsystem. TN on antenna accommodation. FFRF-ASPI-TN September RD[8] Formation Flying RF Subsystem. ICD between Darwin FF RF S/S and GNC S/S. FFRF-ASPI-TN December RD[9] Feasibility Study of the Demonstration Mission (ESA contract no: 17597/03//NL). EADS Austrium. RD[10] Darwin Payload Definition Document (SCI-A/2005/301/Darwin/DMS). ESA June 2005 RD[11] SSTL GPS Receiver SGR05 data sheet: RD[12] GPS antenna GPS S Series. Sensor Systems Inc. USA.. RD[13] Space Flight Technology, German Space Operations Center (GSOC). (Deutsches Zentrum für Luft- und Raumfahrt) DLR. e.v. Phoenix GPS for Proba-2 Interface Control Document. O. Montenbruck, M. Markgraf. Doc. No. : PROBA2-DLR-ICD Version : 0.9. Date : May 19, DLR RD[14] User s Manual for the Phoenix GPS Receiver. O. Montenbruck, M. Markgraf. Doc. No. : GTN-MAN Version : 1.4. Date : May 13, (Deutsches Zentrum für Luftund Raumfahrt) DLR RD[15] High Precision Optical Metrology for DARWIN: Design and Performance, B. Calvel et al., 6 th International Conference on Space Optics, Toulouse 2004 RD[16] TNO-TPD Contributions to High Precision Optical Metrology, a Darwin Metrology Breadboard for ESA, A. Verlaan et al., SPIE meeting, Denver 2004 RD[17] Principe d un coronographe, (M:\_Study\Miscellanous\Coronagraph\Principe d un coronographe.doc)

214 Page 212 of 231 RD[18] Formation flyers applied to solar coronal observations: the ASPICS mission, (M:\_Study\Miscellanous\ASPICS\ SPIE-2005-ASPICS.pdf) RD[19] Yamanaka, K. Ankersen, F. 2002, New State Transfer Matrix for Relative Motion on an Arbitrary Elliptical Orbit. Journal of Guidance, Control and Dynamics. 25 (1), pp RD[20] Spacecraft Structures and Mechanisms, Thomas P. Sarafin, RD[21] Interface Control Document Cryosat Cold Gas Thruster (SV14-001,-002), CS-ID- PAL-RU-0010, S.J. Edwards, Issue1, January 2003 RD[22] Steady-State Thrust Test of Cryosat Cold Gas Flight Unit, MPE/464/LAB, E. Chesta, Issue1 rev 3, 10 February 2004 RD[23] Steady-State Thrust Test of Cryosat OCT Flight Unit Spare S/N 020, TEC- MPE/2004/75/AB-RC-DDC, R.Correia, Issue1, 19 November 2004 RD[24] Cold Gas Feed system for Cryosat- User s manual, CS-MA-PAL-RG-0033, M.Y. Griffith, IssueB, June 2003 RD[25] RD[26] Thrust Verification Tests of the Goce Cold Gas Thruster, TEC-MPE/2004/45/AB, A. Bulit, Issue1, 20 August 2004 RD[27] TCV Thrust and ISP Verification Test Plan / Test Procedure, GO-PR-BRA-0016, J. ELST, 30 June, 2004 RD[28] RD[29] Cold Gas: Technology Status and Plan, PowerPoint Presentation, C. Hunter, 22 June 2005 RD[30] Darwin Payload Definition Document, SCI-A/2005/301/DARWIN/DMS/Lda, L. d Arcio, Issue 1 rev2, 31 August 2005 RD[31] RITA Performance Assessment for DARWIN - Technical Note, DARWIN-1000-TN- 001-EST, D. Feili, Issue1, 17 November 2004 RD[32] EITA, Electron-Bombardment Ion Thruster Assembly, PowerPoint Presentation, N.Wallace, ARTEMIS Ion Propulsion Workshop RD[33] Endurance Tests of 150µN FEEP Microthrusters, IEPC , M. Andrenucci, L. Biagioni, F. Ceccanti, M. Saviozzi, D. Nicolini, November 2005 RD[34] IN-FEEP Characterisation of Microthruster Model, TRP Final Presentation, M.Tajmar, October 2005 RD[35] Technical note on µnrit predevelopment, Technical Note, UNI-GI-DOC , D.Feili, Issue1, 21 February 2005 RD[36] A Miniaturized HET System on board of VEGA-1 P/L, Technical Note, TEC- MP/2005/165/DDC, D. Di Cara, Issue1, 11 September 2005

215 Page 213 of 231 RD[37] Radiofrequency with Magnetic field ion Thruster (RMT): review of the Engineering Phase accomplished under ASI Contract, IEPC-2003-, M. Capacci, G. Matticari, G. E. Noci, A. Severi RD[38] Feasibility Study of the Proba 3 Formation Flying Demonstration Mission, EF.NT.JB , EADS Astrium, Issue 03, RD[39] ECSS-E-50-05A, Radio Frequency And Modulation Standard, published 24 Jan 2003 RD[40] CCSDS B-6. Telemetry Channel Coding Standard. Blue Book. Issue 6. October RD[41] PSS , Ranging Standard, March 1991 RD[42] CCSDS B-5. Packet Telemetry Standard. Blue Book. Issue 5. November RD[43] CCSDS B-3. Telecommand Part 1 Channel Service. Blue Book. Issue 3. June RD[44] CCSDS B-2. Telecommand Part 2.1 Command Operation Procedures. Blue Book. Issue 2. June RD[45] CCSDS B-2. Telecommand Part 3 Data Management Service. Blue Book. Issue 2. June RD[46] ECSS-E-70-41A, Telemetry And Telecommand Packet Utilization, published 30 January 2003 RD[47] CCSDS B-2. Proximity-1 Space Link Protocol Data Link Layer. Blue Book. Issue 2. April 2003 RD[48] CCSDS B-1. Proximity-1 Space Link Protocol Physical Layer. Blue Book. Issue 1. April 2003 RD[49] CCSDS B-1. Proximity-1 Space Link Protocol Coding and Synchronization Sublayer. Blue Book. Issue 1. April RD[50] Formation Flying RF Subsystem. RF Sub-System Architectural Design. 14 December RD[51] Formation Flying RF Subsystem. TN on antenna accommodation. FFRF-ASPI-TN September RD[52] : S and X band patch antenna Surrey Satellite Technology Ltd. data sheet. RD[53] IEEE web site: RD[54] Institute of Electrical and Electronics Engineers, Inc., IEEE Std , IEEE Standard for Information Technology Telecommunications and Information Exchange between Systems Local and Metropolitan Area Networks Specific Requirements Part 15.4: Wireless Medium Access Control (MAC) and Physical Layer (PHY) Specifications for Low Rate Wireless Personal Area Networks (WPANs). New York: IEEE Press RD[55] Current Activities and Status of the ESA-ESTEC Sponsored Wireless Onboard Spacecraft Working Group. DASIA 2004 Wireless Session paper

216 Page 214 of 231 RD[56] ESA Wireless Onboard Spacecraft Web Site : RD[57] Geraldine Artaud, Patrick Plancke, Rodger Magness, Dick Durrant, Chris Plummer. IEEE : Wireless Transducer Networks. DASIA RD[58] Parkinson and Spilker, Eds. Global Positioning System: System and Applications I. Progress in astronautics and aeronautics. Volume 163. RD[59] Formation flying RF S/S Final Report.. Contract number 15511/01/NL/EC. Alcatel Space. RD[60] CAN WG Website: ftp://ftp.estec.esa.nl/pub/ws/wsd/can/can- WG/Documents/Documents.htm RD[61] ESA IP Cores: RD[62] ESA TECHNOLOGY DOSSIER (TOSE-1B-DOS-4) WIRELESS ONBOARD SPACECRAFT AND IN SPACE EXPLORATION (Annex A) - Short-range RF Wireless (Proximity) Networks Technology Assessment for Space Applications RD[63] New alliance formed among Surrey Satellite, U of Surrey Space Centre (UK), Angstrom Aerospace(SE), TPD-TNO and TU Delft (N) regarding implementing COTS-derived RF Wireless for both intra-s/c data handling and inter-s/c formation-flying, miniaturisation and mass reduction in spacecraft. RD[64] All of the above have current and ongoing RF Wireless developments/spacecraft demonstrators. TU Delft Delfi-C3 will launch in (all organisations are represented in our Wireless e-group spacewlan). RD[65] - Ruggedized and Battle Ready Military and US DoD WLAN RD[66] - Military Battlefield WLAN RD[67] - Military WLAN collection of papers RD[68] - Military WLAN list of available HW and specs RD[69] - Military WLAN Antennae RD[70] Note that ESA has recently reiceved a a study made by Alcatel for using COTS for ESA programmes: MILITARY SYSTEM LEVEL products can be used with minor deltaqual! RD[71] - Mars Proximity Telecom (radiation tolerant, wireless TRL network node for telecom and nav.) RD[72] - radiation test for WLAN equipments INSIDE nuclear reactors and plants RD[73] Megarad Wireless communication system RD[74] ftp://ftp-eng.cisco.com/lwood/cleo/readme.html - COTS cisco routers in low earth orbit

217 Page 215 of 231 RD[75] ECSS-E-10-02A Space Engineering Verification RD[76] ECSS-E-10-03A Space Engineering Testing RD[77] ESA bulletin 100 The Integration and Testing of XMM RD[78] ESA Cost Engineering Chart of Services, iss.3, 17 August 2004 RD[79] ECSS-E-10-02A, 17 November 1998 RD[80] TEC-ICE/GRE/HJ/2004/003, TRL Definitions, 01 November 2004 RD[81] TEC-SYC/GRE/SA/2005/021_01, Cost-Risk Analysis Procedure, draft

218 Page 216 of 231 This Page Intentionally Blank

219 Page 217 of ACRONYMS Acronym V AIT AIV AM0 ATB BB BCR BDR BER BOL BP BPSK CBS CCD CCR CDMA CFDP CFI CFRP CLS CMA CME Comm/PV COTS DCR DDS DMM DoD DOF DTCP DWI Definition Velocity increment (delta-v) Assembly Integration & Test Assembly, Integration and Verification Air Mass Zero Avionics Test Bench Breadboard Battery Charge Regulator Battery Discharge Regulator Bit Error Rate Beginning Of Life Black Paint Binary phase shift keying Cost Breakdown Structure Charged Coupled Device Corner-Cube Retro-reflector Code Division Multiple Access CCSDS File Transfer Protocol Call for ideas Carbon Fibre Reinforced Plastic Coarse Lateral Sensor Cost Model Accuracy Corona Mass Ejection Commissioning / Performance Verification Commercial Off The Shelf Dedicated Control Room Data Dispositioning System Design Maturity Margin Depth Of Discharge Degrees of Freedom Daily TeleCommunication Period Dual Wavelength Interferometer

220 Page 218 of 231 Acronym EITA EM EMMA EOL EP EPE EPSA ESTRACK FCT FCU FDR FDS FEEP FER FF FLS FM FOP FOS FTS G/S GaAs GEO GEOS GFCU GFF GNC GNSS GPS GTO HET HEO HGA Definition Electron Ion-bombardment Thruster Assembly Engineering Model Three dimensional Darwin formation End Of Life Electric Propulsion External Project Environment Electric Propulsion Subsystem Assembly ESA Ground Tracking Network Flight Control Team Flow Control Unit Flight Dynamics Room Flight Dynamics Support Field Emission Electric Propulsion Frame Error Rate Formation Flying Fine Lateral Sensor Flight Model Flight Operations Plan Flight Operations Segment Fringe Tracking Sensor Ground Station Gallium Arsenide GEostationary Orbit GEOstationary Satellite Gas feed line & Flow Control Unit Geometrical Fixed Frame Guidance Navigation and Control Global Navigation Satellite System Global Positioning System Geostationary Transfer Orbit Hall Effect Thruster Highly Eliptical Orbit High Gain Antenna

221 Page 219 of 231 Acronym HK HKTM HPCG HPOM HW IA IAR ICC ICDU IF IFP INTEGRAL IOB IP IS ISD ISL J2 L1 L2 LCC LEO LEOP LGA LIT LPF LPF PM LPO LPO LTV LVLH MCR Definition House keeping data Housekeeping Telemetry High Priority Command Generator High-Precision Optical Metrology Hardware Incident Albedo flux Integer Ambiguity Resolution Interferometer Constellation Control Integrated Control and Datamanagement Unit Interface Internal Final Presentation INTErnational Gamma-Ray Astrophysics Laboratory Input Output Board Incident (Infrared) Planetary flux Incident Solar flux Inter-Satellite Distance Inter-Satellite Link Main term of development of central gravitational potential in spherical harmonics corresponding to flattening Libration (or Lagrange) collinear point between primary and secondary body Libration (or Lagrange) collinear point along primary-secondary line ahead of secondary body Life Cycle Cost Low Earth Orbit Launch and Early Orbit Phase Low Gain Antenna Listen-In Test Smart 2 - Lisa Pathfinder Lisa Path Finder Propulsion Module Libration Point Orbit Lagrange Point Orbit the Linear Time Varying Local Vertical Local Horizontal Main Control Room

222 Page 220 of 231 Acronym Definition MCS Mission Control System MGA Medium Gain Antenna MLI Multi Layer Insulation MOC Mission Operations Centre MPPT Maximum Power Point Tracker MPS Mission Planning System NA Neutralizer NNO New Norcia (35m Ground Station) NPU Navigation Processing Unit OBC On Board Computer OBDH On Board Data Handling OBSM On-Board Software Management OCXO Oven-Controlled Crystal Oscillators OPD Optical Path Delay OPD Optical Path Difference P3 PROBA 3 PCDU Power and Conversion Distribution Unit PFM Protoflight Model PISA Principle Investigator Support Area PM Propulsion Module PO Parking Orbit POE Project Owned Events PPCU Power Processing Control Unit PSCU Power Supply Control Unit PSD Power Spectral Density PSR Project Support Room PVA Photovoltaic Array PVAT Position, Velocity, Attitude, Time PXFA Proportional Xenon Feed Assembly QPSK Quadrature Phase Shift Keying R&D Research and Development RAAN Right Ascension of Ascending Node RBF Rotating Body Frame

223 Page 221 of 231 Acronym RF RFDU RFG RFGM RIT RITA RM RMT ROM RVD Rx S/C S/S S2K S3R SA SBC SGM SK SOM SPACON SPF SSMM SSO SSPA SSTL STM SVT TA TC TCS TDMA TJ Definition Radio Frequency Radio Frequency Distribution Unit Radio Frequency Generator Radio Frequency Generator & Matching network Radio-frequency Ion Thruster Radio-frequency Ion Thruster Assembly Reconfiguration Module Radiofrequency with Magnetic field ion Thruster Rough Order of Magnitude Rendezvous and Docking Receiver Spacecraft Sub-System SCOS2000 Sequential Switching Shunt Regulator Solar Array Single Board Computer SafeGuard Memory Station Keeping Satellite Operations Manager Satellite CONtroller Single Point Failure Solid State Mass Memory Sun Synchronous Orbit Solid State Power Amplifier Surrey Satellite Limited Structural-thermal model System Validation Test Thruster Assembly Telecommand Thermal Control Subsystem Time Division Multiple Access Triple Junction

224 Page 222 of 231 Acronym TM TM TPS TRL TRP TT&C TTFG Tx VDA VEGA WP XFCU XMM Definition Telemetry Thruster Module Telemetry Processing System Technology Readiness Level Technology Research Programme Tracking, Telemetry and Command Telemetry & Telecommand Frame Generator Transmitter Vapour Deposit Aluminium Vettor Europeo di Genarazion Avanzata White Paint Xenon Flow Control Unit X-ray Multi-Mirror observatory

225 Page 223 of 231 APPENDIX A - Generic RF Metrology Design and Development Status A-1 RF Metrology Subsystem Development Background Information ESA initiated the development of a breadboard (BB) under the Technology Research Programme (TRP) activity Formation Flying RF Sub-System under ESA contract n 15511/02/NL/EC (RD[6]). The consortium selected for this study was led by Alcatel Space and comprised GMV as sub-contractor. The activity was divided into 2 phases. Alcatel Space also participated in the phase 2 of the study, in the design of the RF transmitter. Phase 1 was focused on the design of the Darwin hexagonal array and Smart-2 (later for Smart-3) RF Metrology subsystem. Phase 2 consisted of the development of a BB with three nodes intended to demonstrate the generic features of the RF Metrology subsystem concept, and the performances based on laboratory experiments and analysis of the results. The design for Smart-2/Smart-3 consisted of a subsystem able to provide the required functionality for 2, 3 and 4 satellites in formation. The equipment selected for is this design, which has been breadboarded and demonstrated in a lab environment. The architecture of the RF metrology unit is illustrated in Figure A 1. I/O bus Antenna set #1 Antenna set #2 RF front-end control Rx RF mute Digital section Tx1 Tx2 Power converter Power bus Antenna set #3 Up-conv Up-conv 10 MHz OCXO Tx module Figure A-1: RF metrology unit architecture The main features of the equipment are: Same design for all the Tx/Rx units RF link in S band, based on GPS receiver heritage. The signal is described in Table A-1 Navigation signal and message transmission and reception to/from all other RF terminals, in all operational phases, using CDMA/TDMA Number of satellites in the formation can be up to 4, also configurable for 2 and 3 Each RF unit is able to accommodate up to 3 sets of 3 antennas (2Rx+1Tx/Rx each) Ranging and angular measurements are calculated based on the measurements of the 3 antennas in one single set TM/TC local data link. 9 kbps for the Master and 3 kbps for Free Flyers Raw data output (ranging, angular measurements and Hz

226 Page 224 of 231 Relative navigation Hz IAR for both ranging and angular measurements (based on carrier phase) is based on dual frequency Operation either in centralised or in distributed mode, or in both Computation of relative position and velocity in Cartesian and polar coordinates, for all satellites in the formation Relative attitude and attitude rate (when applicable) of the satellites expressed in Euler angles (321 sequence). Collision avoidance sensor providing full RF visibility during deployment and nominal phases Hybridisation is required in 2-nodes configuration, with other attitude sensors (for instance, star trackers) to provide full relative attitude and attitude rate restitution with respect to a Master (body) related frame (and, eventually, with respect to an inertial frame). Freq. Band S Frequency MHz 2210 MHz Bandwidth 8 MHz 8 MHz Polarisation RHCP Access CDMA/TDMA Slot sharing configurable Modulation QPSK BPSK Channel I Q I Code type C/A - C/A Code length Chip rate Mcps Mcps Instant. data - 31 Kbps - rate Data coding - No - Min C/N 0 (equivalent) 42 db.hz 48 db.hz 42 db.hz Table A-1: RF metrology signal definition A-2 RF Metrology Development Status The FF RF BB developed within the TRP activity, Formation Flying RF Sub-System under ESA contract n 15511/02/NL/EC was used to demonstrate feasibility and performance of the concept proposed for the Darwin mission and for the Smart-2/Smart-3 demonstrators, (RD[6]). This BB comprises of three RF terminal units and one PC hosting the real-time navigation processing software. Each RF Terminal transmits and receives a TDMA/CDMA L-band GPSlike signal, to and from all other RF Terminals. The BB is set up in a laboratory and the RF Terminal units are connected with cables. Specific devices are used to represent the effect of dynamics, different formations configurations and multipath. Major differences of the BB with the proposed design were the following. The signal transmission occurs in L-band and single frequency, instead of S-band (as required by the SCFG group) and dual frequency. Note that Integer Ambiguity Resolution (IAR) is not possible with a

227 Page 225 of 231 single frequency approach. The data transmission is performed at low data rate (50bps) in one terminal only, for the initial TDMA synchronisation, instead of several Kbps in quadrature to the ranging signal. The navigation processing is running on a PC instead of being embedded in the Tx/Rx unit. The breadboard is illustrated in Figure A-2. Two Rx/Tx units are located on the racks, and the third one is based on an elegant BB, located on the table. Figure A-2: FF RF Metrology BB. View of the test bench. The navigation processing is running on a PC instead of being embedded in the Tx/Rx unit. Two Rx/Tx units are located on the racks (S/C 02 and 3), and the third one is based on an elegant BB, located on the table (S/C 1) The BB developed during the activity was not fully representative of the final design for Darwin, as already mentioned above. It is the intention of ESA to develop an Engineering Model (EM) of the FF RF Metrology subsystem within the current GSTP-4 programme, to overcome the limitations of the BB and to meet the complete Darwin design as defined within the activity. This EM will also be used to validate the feasibility of the critical issues identified during the activity. The architecture and functionality of the EM will be fully representative of the final Flight Model (FM), with the same electrical design, but using commercial components in some parts of the Tx and Rx sub-elements, that will be replaced by space qualified components for the FM. The complete SW will also be embedded in the equipment, including the signal and navigation processing. A-3 Reference Frames For the purpose of the generic RF metrology development, reference frames as described below have been used. During nominal configuration: The state vector is computed in a reference frame attached to the satellite body. In the nominal configuration, the state vector is computed in the Rotating Body Frame (RBF), defined by :