Executive Summary

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Ref: PLATO-T-ASTR-TN-43 Issue: Issue 1 rev 0 Date: 20 Oct 2009

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Title

Assessment Study of the PLAnetary Transits and Oscillations of stars (PLATO) Mission

Executive Summary

Name & Function Date Signature

Prepared by PLATO STUDY TEAM Oct 5th 2009

Verified by

Approved by

SEBASTIEN BOULADE Authorized by Oct 5th 2009 Study manager

Doc type # WP Keywords Summary

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1 PLATO MISSION OBJECTIVES

PLATO mission objectives The PLAnetary Transits and Oscillations of stars (PLATO) mission aims at detecting and characterizing exoplanets by means of their transit signature in front of a very large sample of bright stars, and measuring the seismic oscillations of the parent stars orbited by these planets in order to understand the properties of the exoplanetary systems. The discovery of exoplanets from ground observations makes use of several methods: transits (as proposed for PLATO), radial velocity measurements of the parental stars line-of-sight movements (the first successful method) and gravitational lensing. The success of ground observations has been highly confirmed by the CNES/ESA/European/Brazilian spacecraft CoRoT (Convection, Rotation and planetary Transits). Astroseismology is also performed on CoRoT, although carried out on different targets. The Kepler NASA discovery mission, launched on March 6th 2009, aims also at finding Earth-like planets with the methods of transit and assessing also some astroseismology. Kepler shall observe 100,000 targets of the main sequence of magnitude <14, in a FOV of 105 deg². Both CoRoT and Kepler missions feature limitations in terms of minimum planet size, maximum orbital period, number of detected exoplanets and capability of further characterization of exoplanets and their host stars. PLATO will offer order of magnitude improvement of the science with respect to CoRoT and Kepler missions, filling the need for a further generation mission, observing more stars with increased magnitude and observing significantly smaller exoplanets, with significantly longer orbital periods.

PLATO Science objectives The objective of the PLATO mission is to detect planetary transits in front of stars that can be characterized in terms of fundamental physical parameters. The stars characterization is obtained both from PLATO data themselves via asteroseismology (stellar masses and ages are measured), from the ESA GAIA mission (stellar radius at first order, augmented by PLATO data) and from the ground using e.g. high resolution spectroscopy. When discovered and confirmed, planets characteristics will be inferred from the gathered information on the planet/star radius and mass ratios, coupled to the measurement of the star’s radius and mass. Asteroseismology is therefore a key complement to transits method in the process of planets detection and characterization. In addition to the seismic analysis of planet host stars, which represents the highest priority goal of the mission, asteroseismology of the many other stars present in the field of view will be used to study stellar evolution. The planets to be detected and characterised are of the same type as the Terrestrial planets in the Solar System (Venus, Earth, Mars), orbiting within the inner part of their systems, where they could in principle be the hosts for life (the “Habitable Zone”). Only approximately 1% of the solar-like stars with planets should show Earth-size transits (due to the geometry of the problem). This poor transit probability, associated to the fact that not all stars feature planetary systems, makes the problem a highly statistical one. A large number of stars shall be surveyed in order to get sufficiently information on planets and stars characteristics. The two major requirements of PLATO are: • Observe continuously (minimizing interruptions) 20,000 bright cool dwarf stars during 2 years with 27 ppm/hr photometric accuracy. • Observe continuously (minimizing interruptions) 250,000 faint cool dwarf stars during 2 years with 80 ppm/hr photometric accuracy

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Measured transit

depth

Planet movement on its orbit Light signal received by the observer

transit

No transit

Figure 1-1: The transit method for detection of planets (Left) and astroseismology using amplitude spectra of the signal (Sun oscillations - Right) The method of transits is based on the characterization of the continuous signal received from the star, where an occultation by the planet can be analysed. The geometry of the observation requires a large number of stars to be surveyed to detect planets. The spectrum of stars will be measured on PLATO in the 0.2 µHz-10 mHz range with high accuracy of frequency separations measurements

2 FROM SCIENCE TO SYSTEM REQUIREMENTS

Measurement based on aperture photometry To achieve its scientific objectives, PLATO is relying on the aperture photometry technique: a star is constantly monitored by the telescope and the light collected from this star on the detector is measured by addition of the signal of the pixels in, and at the vicinity, of the star image. The window, called “aperture” is the set of pixels where the signal is collected. The aperture shall best fit the star image on the detector so that to minimize the noise coming from the background or other stars, but shall collect as much as possible signal from the target star, to improve signal to noise ratio. The collected signal for a star is therefore measured along time in a “light curve”. Un-interrupted measurement of this light curve allows detection of a transit in the time domain, or astroseismology science when passing in the frequency domain. There is no absolute photometry here, nor imaging quality and the PLATO main requirements are therefore linked to this measurement principle based on relative photometry only: the telescope shall continuously monitor a high number of stars to obtain un-interrupted light curves, and the star position on the detector in the aperture shall be constant so that the only variations of the light curves do correspond to the science signal.

Performance requirements (number of stars and accuracy) Seven different samples have been defined as targets for observation by PLATO (Figure 2-1 below). These samples are characterized by the number of stars to be observed (cool dwarfs stars mainly, except for sample #3b where the whole main sequence is considered) and the accuracy of the observation, expressed in photonic noise level. The main requirement for non-photonic noise level is that it shall be below 1/3rd of the photonic noise one. Beyond noises induced by the detection chain, the main source of this non-photonic noise is the displacement of the star image inside the aperture defined by the pixels where the signal is summed. Whenever a displacement occurs, the collected signal changes because of several effects such as Pixel Response Non-Uniformity (PRNU), truncature of the image by the aperture, confusion of the signal by a nearby star entering the aperture, or pixels weights factors de-optimisation. The sum of all these effects defines a requirement on the maximum movement of the star image on the detector that can be tolerated

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(this movement comes mainly from the spacecraft attitude control system accuracy, differential velocity aberration and thermo-elastic deformations). Based on the heritage gained on CoRoT, this requirement was assumed to be 0.2 arcsec rms (corresponding to 10 % of stars that would not meet the 1/3rd requirement due to this jitter effect in the magnitude 10-11 range domain). When analysing the relation between the required total collective area of the telescope to meet the photonic noise level, and the required Field of View (FoV) to obtain the proper number of stars with this collective area (by eliminating in the equations the magnitude parameter), it can be demonstrated that the sample #1 is actually the driving requirement (others are met when sample #1 is), and that a whole set of potential designs for PLATO can be positioned on the curve drawing FoV as a function of the collective area. This curve (Figure 2-2) was used extensively during the assessment study to compare potential PLATO designs (high FoV, low collective area or low FoV, higher collective area, FoV splitting, overlapping or duplication, …) and their respective performance with respect to the requirements. This curve was also used to demonstrate analytically that the sample #2b requirement (corresponding to a re- visit of sample#2 stars were a planet has been detected, but with a higher accuracy) would lead to a design featuring less performance for sample #1. For this last reason, the sample #2b objective was not retained for the proposed PLATO reference design, although it could be implemented in further phases.

Sample #1 Sample #2 Sample #2b Sample #3a Sample #3b Sample #4 Sample #5 Number of stars 20,000 80,000 ~ 80 1,000 300 3,000 250,000 Magnitude range NS NS NS Mv ≤ 8 Mv ≤ 8 Mv ≤ 8 (G) 11-14 (G) 27 ppm/h Photonic noise level 27 ppm/h 80 ppm/h 27 ppm/h (G) 27 ppm/h 27 ppm/h 80 ppm/h (G) 2 colours Nb of observed star fields 2 2 NS 2 2 1 2

Figure 2-1: Synthesis of MRD V2.1 requirements on performance (number of stars, accuracy) (NS = Not Specified, G = Goal requirement). Sample #2b was not retained as a requirement for this assessment phase since it would lead to a design featuring less performance on sample #1.

2000

1800 Eddington Kepler 1600

1400 towards larger FOV, 1200 smaller aperture Zone where the design is oversized with respect 1000 to PLATO requirements

FoV (deg²) 800

600 towards smaller FOV, larger aperture 400 Curve of minimum design meeting PLATO Requirements Not Meeting 200 PLATO Requirements 0 0,1 0,2 0,3 0,4 0,5 0,6 0,7 0,8 0,9 1 1,1 1,2 1,3 1,4 1,5 Collecting area (m²) Figure 2-2: Sample #1 requirement curve This curve can be calculated to account for systematic noises (saturation, jitter, …). These noises have been accounted for in the system trade-offs for the instrument design.

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3 PLATO MISSION CONCEPT AND OPERATIONS

The PLATO mission is based on a long uninterrupted high precision photometric monitoring of large samples of stars, in a ‘starer’ concept. The spacecraft is launched from the Kourou space centre with Soyuz-ST and reaches its final eclipse-free orbit around the L2 langrangian point after the transfer phase. On this orbit, the spacecraft will continuously point a given zone of the sky, at high ecliptic latitude, with a large field of view of ~1800 deg². The mission will be divided in three phases. In the first two phases, long monitoring of two successive fields will be performed. A third step&stare phase at the end of the mission will be used to extend the sample of stars surveyed. The duration of the mission is 7.5 years including a one year extension as a goal. All along the continuous observation over a given zone of the sky (at fixed inertial attitude), a sunshield accommodated on the spacecraft protects the instrument from the direct light of the Sun. Every 6 months, the spacecraft is rotated around its pointing axis (line of sight of the instrument) by 180 deg to orientate for the next 6 months its sunshield in the Sun direction. During these 6 months periods, the spacecraft attitude is kept rigorously constant: the solar array of the spacecraft and the communication antenna is rotated once every three days to point respectively at the Sun and at the Earth. Once per month, an orbit control manoeuvre is performed while keeping also the same attitude. This approach ensures that interruptions of science signal collection are minimized.

2 years (goal 3) observation over sky zone 1 2 years (goal 3) observation over sky zone 2 2 years observations in step&stare mode

Lissajoux orbit around L2 Once every 6 months - Free insertion, large amplitude - Avoid eclipses for 6 years

Soyuz launch from Kourou

Operations - ESOC Darmstadt - ESA Cebreros 35 m - X band communications Once every 3 days

Figure 3-1: The PLATO mission

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4 SPACECRAFT ARCHITECTURE

The spacecraft is constituted by the Payload Module (PLM) including 12 telescopes, constituting the whole PLATO instrument, a deployable sunshield (Deployable Sunshield Assembly – DSA) that is stowed at launch, and a Service Module (SVM) providing all resources to the PLM, including in particular a deployable solar array and two High Gain Antennas (HGA, each equipped with a 2-axis mechanism).

Payload Modeule (PLM) : 4 Telescope Assemblies (12 telescopes)

Deployable sunshield (DSA)

PLM Radiator for 1100 electronics

Service Module (SVM)

6970 mm 5200 mm

4400 mm

Figure 4-1: Spacecraft architecture The Solar Array is deployed just after separation of the launcher, during the initial sequence.

5 PAYLOAD MODULE

PLM overall architecture and telescopes design The payload module is constituted by 12-telescopes with two simultaneously observed FoV: each telescope has a very large total FoV (900 deg² for a total ~1800 deg² useful) allowing observation of brighter stars, and the full FoV of the PLATO instrument is available with rotation every 6 months around the LoS. Each telescope is a catoptric 2-mirror telescope entirely manufactured in SiC. Telescopes are grouped in “Telescope Assemblies” (TA) by 3 to optimize the accommodation under the Soyuz fairing and minimize mass. This configuration was selected for its simplicity (simple opto-mechanical design with no corrective lenses), lightweight & low thermal sensitivity advantages, and for the obtained large FoV by rotation around the symmetry axis of the mirror.

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Telescope metering structure: Telescope metering structure: struts M2 support plate

Symmetry to 180 deg rotation every 6 months FP Secondary mirror 6 telescopes sighting FoV #1 Telescope main base-plate M1 6 telescopes sighting FoV #2

Primary mirror

Telescope assembly isostatic mounts Figure 5-1: Telescopes Optical configuration and FOV partition

An important characteristic of this design is also its scaling capability without affecting optical performances: the size of the pupil can then by changed without any impact on the optical performances, offering a valuable robustness and flexibility of the system in terms of mass. The thermal control of the instrument is fully passive, with no risk of perturbation on the PLATO science measurements, taking maximum advantage of the PLATO naturally stable thermal environment. The obtained CCD temperature is below 130K. To protect the payload from the Sun, a deployable sunshield has been designed, derived from the GAIA deployable sunshield, hold at launch with a strap that is released once in orbit and kept safe in a winding system. The focal planes design consists of 176 CCDs in total and includes: • 14 baseline CCD arrays per telescope. Each CCD is a large size one (maximum size to fit two arrays in 6” wafer) with large pixel pitch (29 µm). This result in a moderate format of 5.4 Mpixels per CCD. • Two “bright-star mode” CCDs (per telescope, one for each colour) in 4 telescopes for Sample #3b coloured measurements, with the same design than the baseline ones but operated in framestore mode with a reduced integration time. This approach allowed to use the sample #1 telescopes for sample #3b , without need of additional telescopes. • The detection chain, with two video chains per CCD, a pixel rate compatible with GAIA16-bit ADC, and one interconnection module (I2M) per focal plane to group the interfaces into a single SpaceWire serial data link to the payload data processing units hosted on the SVM. The I2M also features a “windowing” FPGA that collects only the useful signal in the image (aperture, imagettes) and send it to the processing units.

Payload data processing design The payload data processing is composed by: • Five Digital Processing Units (DPUs, one per Telescope Assembly + 1 redundant), in charge of processing the pixels of the masks to obtain light curves and centroids, of processing the CCDs data (background, imagettes, etc.) and of the overall instrument control function. They include all the required mass memory devices to store science data before transmission to ground. • One Interconnect Unit (ICU), in charge of routing the science signal from the telescopes to the DPUs, and from the DPUs to the spacecraft data handling and communication system (for transfer of science data to ground). This ICU is not the Instrument Control Units (ICU) of the previous CDF report but here simply a router of all signal so that the instrument control function can be handled by the DPUs themselves.

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Figure 5-2: PLATO PLM configuration The electronics are attached to the stable structure that supports the 4 telescopes assemblies 1 Silica plate protects the focal planes from direct radiations in the FOV.

CCD readout and Geometrical FoV flexprint

Full-performance FoV

Baseline mode CCD operated in full frame Bright star CCD mode (14 off) operated in framestore FoV = 98 deg² mode (2 off)

Figure 5-3: Payload focal plane accommodation

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6 SERVICE MODULE

The Service Module (SVM) architecture proposed for PLATO is a cost efficient architecture based on off- the-shelf components with strong re-use of GAIA (among which AOCS sensors and actuators, data handling system, power control, and overall structure architecture and parts). Specific tailoring is performed for the instrument interfaces and the communication subsystem. All spacecraft command and control tasks are performed by the On-board Computer (OBC) featuring internally redundant processor modules with built-in data memorisation capability for housekeeping storage and mission management purposes. The Data Handling is based on packet telemetry and telecommand (TM/TC). The OBC interfaces for TM/TC with the payload via a dedicated MIL-STD 1553B bus, and to other spacecraft units via another MIL-STD-1553B bus. A mission customised Electrical Interface Unit (EIU) connected by SpaceWire links to the OBC gathers all other input/output interfaces with the rest of the spacecraft. The power subsystem is based on a deployable and one axis rotating solar generator, followed by Maximum Power Point Tracking (MPPT) regulation towards a regulated 28V bus. The communication subsystem includes two transponders with hot redundant receivers and cold redundant transmitters. Omni-spherical coverage is achieved by two low gain antennas (LGA) with hemispherical coverage. The payload data is amplified by amplifiers (TWTAs) and transferred to the ground station by one of two steerable high gain antennas (only one used at a time, depending on the spacecraft orientation). The TWTAs are also used for TM transfer via the low gain antennas. The PLM will be installed on top of the service module in a thermal and mechanical stable manner. The primary structure is optimized for carrying a relatively heavy payload (significantly higher mass as for GAIA). The thermal design is based on passive regulation. The Attitude and Orbit Control System (AOCS) concept uses heritage from GAIA (the developed Micro-Propulsion System - MPS) and offers high accuracy pointing, largely meeting science observation requirements. In particular, should the stability requirements (0.2 arcsec jitter requirement for 10% corresponding impact on sample #1 stars) become more stringent in potential future phases, the high accuracy AOCS designed during this assessment study would be a key asset of the proposed configuration.

Figure 6-1: Spacecraft functional architecture

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7 PLATO PERFORMANCES AND BUDGETS

The Proposed PLATO spacecraft design meets performance requirements with good margins, with more than 22,000 cool dwarfs observed at 27 ppm/hr, more than 260,000 stars observed at 80 ppm/hr, and a large FoV (1800 deg²) enabling the observation of bright stars (Mv < 10.5 in Sample #1).

Star samples Instrument level: over 1h & all telescopes Sample #3b Sample #1 Sample #2 Sample #3a Sample #4 Sample #5 (each colour) Requirements Photonic noise (ppm/h) 27 80 27 27 27 80 Required number of stars in sample 20000 80000 1000 300 3000 250000 Number of observation periods 2 2 2 2 1 2 Required limit magnitude in sample NS NS 8 8 8 (G) 14 (G) Required number of stars in telescope FoV 10797 52280 500 153 3239 163375 Performances: number of stars & magnitude range Min magnitude set by saturation 7.94 7.94 7.94 5.49 7.94 7.94 Limit magnitude to meet number of stars 10.27 11.63 8.40 7.70 9.24 12.60 Limit magnitude to meet photometry 10.36 12.72 10.36 7.70 10.36 12.72 Number of observed stars 23885 374181 1000 305 11942 374181 Margin for confused stars 2.7% 30% 0% 0% 2.7% 30% Margin for saturated stars 5.3% 0.7% 1.7% 5.3% 0.7% Number of expected valid stars 22014 260093 1000 300 11007 260093 Figure 7-1 : Star count & magnitude range performances Sample #3a exhibits no margin since star magnitude is the limiting factor: the requirement on magnitude is not met with 8.4 for a requirement of 8, but this no-compliance was considered acceptable. 92 stars are observed below mv 8.

The proposed PLATO spacecraft design meets mission requirements, compliant within Soyuz mass & volume with 20% system margin. The Plato launch mass is 2050 kg. The maximum power used in science observations with data downlink is 1500 W. Finally, the proposed design features valuable robustness versus potential mass evolutions through PLM moderate scaling and stability requirements evolutions through high accuracy pointing.

Plato PLM and SVM masses SVM System mass 744 kg

PLM System mass 1174 kg

Propellant Hydrazine budget 76 kg Cold Gas propellant budget 56 kg Launch mass Spacecraft Launch mass 2050 kg Launcher adapter 90 kg Total Launch mass 2140 kg Figure 7-2 : PLATO mass system budget

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8 PLATO DEVELOPMENT PLAN

Phase A/B1 is assumed to start at the mid of 2010 after completion of the Cosmic Vision selection process in the second half of 2009. PLM & SVM are assumed to be delivered to Prime contractor at the end of 2016 for spacecraft integration & tests. Plato development schedule is compatible with a launch in 2018, including 3 months AIT margins, 6 months ESA margins, and a 6 months margin for ESA down-selection of M-class missions.

PLATO Mission Schedule2010 2011 2012 2013 2014 2015 2016 2017 2018 1Q10 2Q10 3Q10 4Q10 1Q11 2Q11 3Q11 4Q11 1Q12 2Q12 3Q12 4Q12 1Q13 2Q13 3Q13 4Q13 1Q14 2Q14 3Q14 4Q14 1Q15 2Q15 3Q15 4Q15 1Q16 2Q16 3Q16 4Q16 1Q17 2Q17 3Q17 4Q171Q182Q183Q184Q18

Project phases SRR PDR CDR FAR System Definition Study System Phase B System Phase C/D Final down-selection and ITT process

PLM TDAs CCD TDA Sunshield TDA PLM development & AIT Focal Planes Telescope Assemblies Telescopes integration PDP Sunshield

SVM development & AIT

Spacecraft AIT AIT Margin ESA Margin Launch campaign Figure 8-1: Spacecraft schedule

ACRONYMS

AOCS Attitude and Orbit Control System OBC On-Board Computer BDR/BCR Battery Discharge / Charge Regulator PDP Payload Data Processing CCD Charge Coupled Device PFM Proto-Flight Model CDF Concurrent Design Facility PI Principal Investigator DPU Data Processing Unit PIP Payload Interface Plate DSA Deployable Sunshield Assembly PLM PayLoad Module EIU Electrical Interface Unit PSD Power Spectral Density ESA European Space Agency PSF Point Spread Function FCL Fold-back Current Limiter. SciRD Science Requirement Document FEE Front-End Electronics SM Structural Model FM Flight Model SNR Signal To Noise ratio FPA Focal Plane Assembly SVM Service Module HDRM Hold-Down & Release Mechanism TA Telescope Assembly I2M Interface & Interconnection Module TBC To Be Confirmed LCL Latching Current Limiter TDA Technology Development Activity LGA Low Gain Antenna TM/TC Telemetry / telecommand LOS Line Of Sight TWTA Travelling Wave Tube Amplifier MOC Mission Operation Centre UART Universal Asynchronous Receiver/Transmitter MPPT Maximum Power Point Tracking MPS Micro-Propulsion System MRD Mission Requirement Document