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    The Euclid mission architecture is strongly driven by the science requirements and programmatic constraints. Major
    issues of the sky survey are its speed, depth, precision, and imaging quality while the main programmatic constraint is
    the mission duration. The survey speed is guaranteed by the combination of a large field of view, about 0.54 deg2
    , and an
    optimised survey strategy. Ensuring the high image quality leads to demanding requirements on the pointing and thermoelastic
    stability. The survey depth leads to a minimum telescope aperture, dedicated baffling design, low temperature
    optics and detectors, a cold telescope for low near-infrared background and on-board data processing for the noise
    reduction of the near-infrared detectors.
    A large amplitude orbit around the second Sun-Earth Lagrange point (SEL2) has been selected because it imposes
    minimum constraints on the observations and allows scanning of the sky outside the galactic latitude b ±30 degree band
    around the Milky Way within the mission duration. The Euclid spacecraft will be launched from the Guiana Space
    Centre, Kourou, on board a Soyuz ST 2.1-B. The launch date and time determine the ellipticity and size of the
    operational orbit and influence the Sun-Spacecraft-Earth angle plus the daily visibility from the ground station. The
    launch is possible in most of the days of the year with minor restrictions to avoid eclipses during transfer and in the
    operational orbit, and by the angle between the Sun and telescope aperture during transfer. Once in operational orbit the
    spacecraft performs a step-and-stare scanning of the sky.
    2.1 Survey design
    Euclid aims to cover very large areas with great stability, thermal stresses must be minimised and this impacts on
    operations. In fact, the allowed pointing range is limited to observe orthogonally to the Sun, in a range of -3º towards to
    +10º from the orthogonal. In practice Euclid can scan parts of a circle on the sky along the ecliptic meridian; the
    visibility at ecliptic equator is ~ one week per semester (the target can be seen six months later along the same circle).
    This visibility period increases with the distance from the equator, up to two small circles at the ecliptic poles, which
    have perennial visibility.
    The elementary observation sequence of a field is composed of four frames of the 0.54 deg2 common area, observed with
    a dither step in-between. During each frame the visual instrument (VIS) and the Near Infrared Spectro-Photometer
    (NISP) spectrometer carry out exposures of the sky simultaneously. Subsequently, because of the disturbing vibration
    from filter wheel rotation, VIS closes its shutter during the remaining exposures while NISP photometric imaging is
    performed. At the end of the last frame, a slew towards the next field is performed. A significant part of the mission is
    ECSURV Summary of Survey(s) Status and perspectives
    Ref:
    Version 1.0
    Date 29/03/2016
    Page 7/19
    3. Wide Survey
    Because of the large amount in number, time and repeats of calibrations, the strategy has to fulfil calibrations
    first, then use the remaining time to observe the wide areas. At present there is a single standard sequence
    for each wide field, as given in the figure (recall there is no longer room for blue grim exposures in the wide).
    The wide coverage achieved year by year can be easily seen in the picture below [ECTile, J. Dinis],
    where horizontal bars show what is observed in a given year (two bars for each year, since one can observe
    either in the leading or trailing direction wrt the orbit). In each bar in the lower panel the smaller dark areas
    are the times reserved for calibrations (J. Amiaux, I. Tereno), while the wider ones are color coded with the
    corresponding regions of the wide on the sky.
    The presented document is Proprietary information of the Euclid Consortium.
    This document shall be used and disclosed by the receiving Party and its related entities (e.g. contractors and subcontractors) only for the purposes
    of fulfilling the receiving Party's responsibilities under the Euclid Project and that the identified and marked technical data shall not be disclosed or
    retransferred to any other entity without prior written permission of the document preparer.
    Current&state&(example&survey)&
    BeYer&space3;me&coherence&
    Much&lower&presence&of&bubbles&
    (s;ll&work3in3progress)&
    How:&&added&patch&adjacency&to&p atch&scoring&
    &&&&&&&&&&&backward3;ling&(backw ard&in&;me)&
    &&&&&&&&&&&patch&compression&
    Mission Operation Concept
    Document part B:
    Reference Survey
    Ref.
    Version:
    Date:
    Page:
    EUCL-EC-RP-8-001
    6.8
    11/09/2015
    48/88
    The presented document is Proprietary information of the Euclid Consortium.
    This document shall be used and disclosed by the receiving Party and its related entities (e.g. contractors and subcontractors) only for the purposes
    of fulfilling the receiving Party's responsibilities under the Euclid Project and that the identified and marked technical data shall not be disclosed or
    retransferred to any other entity without prior written permission of the document preparer.
    o The Solar Panel Solar Aspect Angle (SPSAA) is defined as the angle between the
    spacecraft +XSC axis and the direction to the centre of the solar disk.
    o not exceed variations of the Solar Panel Solar Aspect Angle (SPAA) of up to 10 (TBC)
    degrees
    5.1.1. Elementary Observation Implementation
    The elementary observation sequence over a field is composed of four frames observed with a dither
    step in between. At the end of the last frame, the FoV slews towards either:
    x The next Survey Field of View
    x A calibration field (frequency see calibration sequence). The calibration field can be a high
    density star field located in the Galactic plane or the current observed field, observed stabilised
    with calibration source on for flat field calibration (see VIS and NISP IOCD [AD8] and [AD9]).
    The Field of View duration is (without margin on integration time):
    • 4 x 973 s + 3 x 60 s + 290 s = 4072 s + 290 s = 4362 s
    The exposure time (including read out overheads) in the VIS and NISP are given in VIS and NISP
    Performance report ([AD6] and [AD7]) based on current Space Segment definition (see [AD1]):
    x VIS exposure time = 565 s
    x NISP Spectroscopy exposure time = 565 s
    x NISP Photometry exposure time:
    o Y = 121 s
    o J = 116 s
    o H = 81 s
    Figure 5-4: Nominal Field Observation Sequence.
    To Next Field Dither 01 Dither 02 Dither 03 Dither 04 Δt < 290 s Dither Step 01 Dither Step 02 Dither Step 03 Slew
    60 s 60 s 60 s 290 s
    VIS
    NISP
    Shutter
    10 s
    NISP
    565 s
    VIS
    565 s
    Shutter
    10 s
    FWA
    10 s
    Y
    121 s
    FWA
    10 s
    J
    116 s
    FWA
    10 s
    H
    81 s
    GWA
    10 s
    Dither = 973 s
    Stab
    10 s
    Stab
    10 s
    Stab
    10 s
    Nominal Science Observation Sequence = 4362 s
    Figure 1 Euclid standard observing sequence
    devoted to instrument calibrations and sample characterisation. The former class needs specific observations, ~6 months
    in total, on given targets (spectro-photometric standards, repeated fields for stability and flats). The latter needs repeated
    observation for depth and have different dispersion angles for the same objects or observations on well-known
    astronomical fields, ~6 months in total. Interleaved with calibrations, much time will be devoted to the Euclid Deep
    Fields (EDFs), which will be two magnitudes deeper than the wide survey and cover a minimum total area of 40 deg2
    .
    Because of the need of repeated observations, long observability plays a key role here and therefore possible locations
    are forced to be close to the ecliptic poles.
    The wide area has to solve a number of demanding constraints related to the visibility: the interspersed set of
    calibrations, the zodiacal background (which increases by factors in going from ecliptic poles to the ecliptic equator and,
    moreover, is also time dependent), galactic dust extinction and scattered light. The resulting Euclid sky coverage has to
    exclude the ecliptic plane and the galactic plane and bulge.
    Even though there are many solutions for the sky coverage, it is not obvious to find a solution satisfying all visibility,
    operational, and programmatic constraints. An optimal solution was found for the MPDR by the Euclid Consortium
    Survey group using a novel software, ECTile, developed exclusively for this purpose. The algorithm takes into account
    all the constraints and optimises possible sequences of pointing’s on the sky. In practice the time reserved for
    calibrations is set as an input and the best coverage is found by using the unallocated days to observe the not yet covered
    sky, weighted with a merit function.
    The solution, shown in Figure 2 with colour coding according to the epochs, is close to the theoretically maximum
    achievable (a straight line in the second plot). Towards the end of the survey most of the visible sky has been observed
    previously, so that no new areas have very limited visibility, causing only small increases in the growth curve.
    Figure 2 Left panel: area covered by the wide survey (ecliptic coordinates, colour coding follows the epoch of observation).
    The empty regions reflect the ecliptic equator and the galaxy plane Right panel: growth curve, the increase of the area
    covered by the wide survey as a function of time.
    It must be recalled that the above reference survey is a proof of feasibility, the final survey will be delivered after launch
    and in-orbit performance verification.
    3. SPACECRAFT DESIGN
    The spacecraft can be subdivided in three main parts: a Service Module, a Payload Module, including the telescope, and
    the Scientific Instruments. They are separately described in the following sections.
    3.1 Service Module
    The Service Module (SVM) comprises the spacecraft subsystems supporting the payload operation, hosts the warm
    electronics of the payload, and provides structural interfaces to the Payload Module (PLM) and the launch vehicle. The
    Sunshield, part of the SVM, protects the PLM from illumination by the sun and supports the photovoltaic assembly
    supplying electrical power to the spacecraft. The overall spacecraft envelope, compatible with the Soyuz ST fairing, fits
    within a diameter of 3.74 m and a height of 4.8 m, see Figure 3.
    ECSURV Summary of Survey(s) Status and perspectives
    Ref:
    Version 1.0
    Date 29/03/2016
    Page 7/19
    3. Wide Survey
    Because of the large amount in number, time and repeats of calibrations, the strategy has to fulfil calibrations
    first, then use the remaining time to observe the wide areas. At present there is a single standard sequence
    for each wide field, as given in the figure (recall there is no longer room for blue grim exposures in the wide).
    The wide coverage achieved year by year can be easily seen in the picture below [ECTile, J. Dinis],
    where horizontal bars show what is observed in a given year (two bars for each year, since one can observe
    either in the leading or trailing direction wrt the orbit). In each bar in the lower panel the smaller dark areas
    are the times reserved for calibrations (J. Amiaux, I. Tereno), while the wider ones are color coded with the
    corresponding regions of the wide on the sky.
    The presented document is Proprietary information of the Euclid Consortium.
    This document shall be used and disclosed by the receiving Party and its related entities (e.g. contractors and subcontractors) only for the purposes
    of fulfilling the receiving Party's responsibilities under the Euclid Project and that the identified and marked technical data shall not be disclosed or
    retransferred to any other entity without prior written permission of the document preparer.
    Current&state&(example&survey)&
    BeYer&space3;me&coherence&
    Much&lower&presence&of&bubbles&
    (s;ll&work3in3progress)&
    How:&&added&patch&adjacency&to&p atch&scoring&
    &&&&&&&&&&&backward3;ling&(backw ard&in&;me)&
    &&&&&&&&&&&patch&compression&
    Mission Operation Concept
    Document part B:
    Reference Survey
    Ref.
    Version:
    Date:
    Page:
    EUCL-EC-RP-8-001
    6.8
    11/09/2015
    48/88
    The presented document is Proprietary information of the Euclid Consortium.
    This document shall be used and disclosed by the receiving Party and its related entities (e.g. contractors and subcontractors) only for the purposes
    of fulfilling the receiving Party's responsibilities under the Euclid Project and that the identified and marked technical data shall not be disclosed or
    retransferred to any other entity without prior written permission of the document preparer.
    o The Solar Panel Solar Aspect Angle (SPSAA) is defined as the angle between the
    spacecraft +XSC axis and the direction to the centre of the solar disk.
    o not exceed variations of the Solar Panel Solar Aspect Angle (SPAA) of up to 10 (TBC)
    degrees
    5.1.1. Elementary Observation Implementation
    The elementary observation sequence over a field is composed of four frames observed with a dither
    step in between. At the end of the last frame, the FoV slews towards either:
    x The next Survey Field of View
    x A calibration field (frequency see calibration sequence). The calibration field can be a high
    density star field located in the Galactic plane or the current observed field, observed stabilised
    with calibration source on for flat field calibration (see VIS and NISP IOCD [AD8] and [AD9]).
    The Field of View duration is (without margin on integration time):
    • 4 x 973 s + 3 x 60 s + 290 s = 4072 s + 290 s = 4362 s
    The exposure time (including read out overheads) in the VIS and NISP are given in VIS and NISP
    Performance report ([AD6] and [AD7]) based on current Space Segment definition (see [AD1]):
    x VIS exposure time = 565 s
    x NISP Spectroscopy exposure time = 565 s
    x NISP Photometry exposure time:
    o Y = 121 s
    o J = 116 s
    o H = 81 s
    Figure 5-4: Nominal Field Observation Sequence.
    To Next Field Dither 01 Dither 02 Dither 03 Dither 04 Δt < 290 s Dither Step 01 Dither Step 02 Dither Step 03 Slew
    60 s 60 s 60 s 290 s
    VIS
    NISP
    Shutter
    10 s
    NISP
    565 s
    VIS
    565 s
    Shutter
    10 s
    FWA
    10 s
    Y
    121 s
    FWA
    10 s
    J
    116 s
    FWA
    10 s
    H
    81 s
    GWA
    10 s
    Dither = 973 s
    Stab
    10 s
    Stab
    10 s
    Stab
    10 s
    Nominal Science Observation Sequence = 4362 s
    ECSURV Summary of Survey(s) Status and perspectives
    Ref:
    Version 1.0
    Date 29/03/2016
    Page 8/19
    This is a result of a further optimisation
    wrt the reference for mission PDR,
    shows from the other figure that with
    4400 sec exposures we are able to
    cover the wanted area in the wanted
    time (all the calibrations and EDFS are
    incorporated, as well as the 1 day per
    month of orbit maintenance). The lower
    curve refers to the non observed fraction
    because of lack of visibility of new areas
    at that particular epoch (of course one
    will use those dead times -if any- wisely,
    e.g. reobserve some fields, but at
    present these are too variable and
    sparse to have now a precise plan for
    them).
    However, this fantastic result, close
    to theoretical achievable maximum, can
    only worsen in practice because of a
    number of factors. There is a science issue related to the coupling of visibility vs “good” sky areas: because
    of the choice of fixed exposures, the S/N will vary quite a bit in the Euclid fields, according to the local
    background. This is due to the zodiacal effect (mainly a function of the ecliptic latitude but also time
    dependent) but also to the straylight from the plane of the galaxy and other star rich areas. Therefore we can
    observe 15000 sq degs but are all of them useful? The answer is no but this has to be quantified better by
    the end to end study but it leaves open an interesting possibilities:
    WIDE Possibility #1: see if an increase
    in overlap fraction between nearby fields
    would improve results (e.g. photometric
    dispersion) at some cost in covered area/
    time.
    WIDE Possibility #2: to have a
    quantised approach to in part compensate
    for the different levels of S/N given by fixed
    exposure time and varying backgrounds. I.e.
    to divide the whole areas in blocks to be
    observed with, say two or three different
    approaches in which the whole sequence or
    only VIS/NISP spectral exposure times are
    varied with respect to a fixed global one. The
    figure reflects the goal of a continuous
    variation of T_exp so to have S/N~constant,
    which is not possible.
    A quantised version would still aim to
    decrease the strong gradient expected for an
    global T_exp but the tradeoff with the
    needed increase of calibrations must be
    assessed.
    WIDE Possibility #3: instead of
    observing 15,000 sq degs once, observe
    (15-M)x 103 sq degs once and observe twice
    (M/2)x 103 sq degs or trice (M/3) x 103 sq
    degs so to have two (or three) passes in
    selected areas: this would on only give
    confidence on the overall results and their
    repeatability (repeatability is currently limited
    to a few tens of sq degs in the basic
    reference) but also increase the S/N in those
    areas. This approach therefore can be used
    to observe areas which are interesting from a
    general point of view because have other
    The presented document is Proprietary information of the Euclid Consortium.
    This document shall be used and disclosed by the receiving Party and its related entities (e.g. contractors and subcontractors) only for the purposes
    of fulfilling the receiving Party's responsibilities under the Euclid Project and that the identified and marked technical data shall not be disclosed or
    retransferred to any other entity without prior written permission of the document preparer.
    Weirdo W#N
    first/second year: go faster,
    last: go slower
    Low
    |β|
    fixed
    T_exp
    variable
    T_exp 30
    35 32 28 25
    29 31 30
    High
    |β|
    tech time same, t_exp varies (but keep still in backgr noise limit)

    (

    ))
    ))

    for fixed S/N:
    T_exp ~ Zodi
    Not allowed by instrument teams
    (calibs, stability, CTI reconstr.)
    Low
    |β|
    std exp
    expose 2x, 3x,
    Weirdo W#N+2
    latest epochs: less area but with better sampling and in selected
    areas already observed from the ground (some close to the ecliptic
    plane). Better for legacy and other Xchecks
    Replace worst 1000 sq deg with 300-500
    (double or triple exposures because of
    larger backgr. on “good/interesting” areas
    (some cost to FoM but better science
    overall - ?large slews?)
    Current&state&–&performance&pl ot&
    Survey&performance&improved&a&bi t&(lower&idle3;me)&
    Figure 3 Euclid spacecraft overview
    Mechanical and Thermal Architecture
    The SVM (Figure 4) is an irregular hexagonal base built around a central cone that provides the interfaces with the
    launcher and with the PLM and encloses the Hydr
    Figure 4 SVM overview (the central cone aperture is sealed by MLI belonging to the SVM, not shown here
    Figure 5 SVM equipment accommodation
    The thermal control is based on a passive design using radiators, multilayer insulation (MLI) and heaters operated in
    Pulse Width Modulation. The design drivers are the short-term temperature stability of the PLM conductive and radiative
    interface under the maximum commanded Solar Aspect Angle (SAA) change, and minimal (<25 mW) heat flux into the
    coldest NISP radiator. High performance Kapton MLI is installed on the on SVM top floor, PLM bottom and Sun Shield
    rear side to minimise the heat flux and thermal disturbances onto the PLM.
    Electrical and Data Handling Architecture
    The spacecraft provides 28V regulated power to equipment and instruments electronics through protected lines
    individually commandable provided by the on-board power conditioning and distribution unit (PCDU). The PCDU also
    provides power to the heaters, to the pyro actuators and controls the charge and discharge of the battery. The battery is
    used only during the launch phase and is design to provide up to 419 W of power and a total energy of 300 Whr. In the
    other phases of the mission the sun shield three panels provide a power between 2430 and 1780 Watt depending from the
    spacecraft orientation and ageing of the panels. See budgets in Table 2.
    One centralised on-board computer (Command and Data Management Unit, CDMU) provides spacecraft and AOCS
    command, control and data processing. The CDMU is a modular unit including standard core boards plus dedicated I/O
    boards to interface AOCS and spacecraft units and devices. The Processor Module is based on a general-purpose space
    qualified microprocessor (LEON-FT) with minimum computational power of 40 MIPS and 5 MFLOPS. Two processor
    modules are comprised in a single-failure tolerant unit.
    The number of scientific exposures and high-resolution images generate a high science data volume and require large on
    board memory capable of hosting the 850 Gbit of daily generated. The on-board Mass Memory Unit (MMU) has a
    capacity of 4Tbit EoL sufficient to store 72 hrs of scientific data and 20 days of spacecraft housekeeping. The MMU
    stores instruments data and housekeeping and other ancillary data in named files organized in a two level folders’
    structure.
    The commands and telemetries are distributed and collected mostly via two standard Mil-Std-1553 buses, one dedicated
    to the spacecraft equipment and another to the instruments and mass memory, although some spacecraft equipment have
    dedicated connections. The instruments deliver high volume scientific data via high speed SpaceWire links directly into
    the mass memory. The platform bus handles non-packet remote terminals (RT) and the FGS, and is characterised by
    cyclic communication frames at 10 Hz, linked to the AOCS control cycle. The transfer layer protocol of the science bus
    is based on a cyclical communication frame at 60 Hz, maximising the efficiency of data transfer per communication
    frame.
    Figure 6 CDMS interfaces
    Files stored in the mass memory are downloaded using the standard CCSDS File Delivery Protocol (CFDP) using the
    reliable transfer with acknowledges for the downlink and the simple unreliable transfer for uplink. Both the X and K
    band communication link can be used for the file transfer. The baseline configuration expect the directives of the CFDP
    to be transmitted via X-band to ensure visibility of the file downlink in progress by Ground also in case of adverse
    weather conditions, while data are downloaded via K-band to maximise the data rate. Any other combination is however
    possible.
    Telecommunications
    The telecommunications architecture includes two independent sections: an X-band section used for telecommands,
    monitoring and ranging and a K-band section dedicated to high rate telemetry (Figure 7).
    The X-band section supports uplink of telecommands at two different rates (4 kbit/s and 16 kbit/s), downlink of real time
    housekeeping TM at two different information rates (2 kbit/s and 26 kbit/s), and standard ranging. The X-band section
    uses two X-band transponders, with receivers operated in hot redundancy and cross-coupled with CDMU TC decoders
    and transmitters operated in cold redundancy and cross-coupled with CDMU TM encoders. Three X-band LGA’s with
    hemispherical coverage are used. Two of them (LGA-1 and -2), placed on opposite sides of the spacecraft and working
    in opposite circular polarisations, provide the omnidirectional coverage. LGA-3, mounted to the HGA support structure
    and sharing its pointing mechanism, supports high rate telecommand. The receivers are cross-coupled with LGA-1 or
    LGA-3, and LGA-2.
    Figure 7 TT&C block diagrams, X-band (top) and K-band (bottom)
    The K-Band section supports downlink of recorded science and housekeeping telemetry at two different data rates:
    nominal at 73.85 Mbit/s and reduced at 36.92 M
    between different sky zones). After each slew manoeuvre the wheels are controlled to slow down until friction stops
    them. Keeping the reaction wheels at rest during operation ensures noise-free science exposures by eliminating the
    micro-vibration associated to reaction wheel actuation.
    The micro-propulsion employed for fine attitude control is based on cold-gas Nitrogen thrusters in a configuration of two
    branches with six thrusters each. Four high-pressure tanks provide storage of 70 kg Nitrogen, sufficient for 7 years
    operation with nearly 100% margin.
    Orbit control and attitude control in non-science modes are actuated by two redundant branches of ten 20N hydrazine
    thrusters. In each branch, two thrusters, one on either side of the spacecraft, provide torque-free thrust for the Trajectory
    Control Manoeuvres on the way to SEL2, monthly Station Keeping Manoeuvres at SEL2, and disposal at end of life. The
    other eight thrusters provide force-free torques for angular momentum and attitude control in non-science modes.
    Hydrazine storage is provided by one central tank with 137.5 kg propellant mass capacity with 10% volume margin over
    and above the prescribed delta-v margins.
    System budgets
    The mass budget in Table 1 shows the breakdown at module level based on detailed estimates validated by the
    subsystem suppliers. A lightweight adapter of the same design used on Gaia is employed.
    Table 1 System mass budget: the SVM and the PLM masses include the instrument units located in each module, reported
    below each item
    Current mass [kg]
    SVM 920.6
    Payload warm units 98.0
    PLM 848.3
    Payload cold units 193.0
    Dry mass reserve 113.9
    Propellant 199.2
    Launch vehicle adapter 78.0
    Total launch mass 2160.0
    Table 2 System power budget. The sizing case is shown (maximum payload power demand, communications on;
    end of life; maximum voltage).
    Power [W]
    SVM (@ 28 VDC) 774
    PLM (@ 28 VDC) 88
    Instruments (@ 28 VDC) 392
    System losses (3%) 38
    System margin (20%) 258
    Array power demand (at PCDU input and @
    29.5 VDC)
    1360
    Table 3 RF link budget. Cebreros ground station, 1.77 million km, elevation above horizon 5° (X band) and 20° (K band).
    For the LGA items, degrees in parentheses indicate angle from centre of pattern.
    Antenna Bit
    rate
    [kbit/s]
    Nom.
    Margin
    [dB]
    X band
    TC (90°) LGA1/2 4.0 4.1
    TC (5°) LGA3 16.0 7.3
    TM (90°) LGA1/2 2.0 5.8
    TM (5°) LGA3 26.0 8.7
    K band
    TM (low rate) HGA 73.85 4.8
    TM (high rate) HGA 36.93 7.6
    Table 4 Pointing budgets (99.7% confidence level)
    Requirement
    [arcsec]
    Performance
    [arcsec]
    APE (X/Y axis) 7.5 6.25
    APE (Z axis) 22.5 12.27
    RPE (X/Y axis) 75⋅10-3 70.5⋅10-3
    RPE (Z axis) 1.5 0.275
    3.2 Payload Module
    The Euclid PLM is designed around a three mirrors anastigmatic Korsch Silicon Carbide (SiC) telescope feeding the two
    instruments, VIS and NISP, see schematic in Figure 8. The light separation between the two instruments is performed by
    a dichroic plate located at exit pupil of the telescope. The PLM is in charge of providing mechanical and thermal
    interfaces to the instruments (radiating areas and heating lines). Whereas NISP is a stand-alone instrument with interface
    bipods, VIS is delivered in several separate parts: a focal plane assembly (FPA) connected to proximity electronics,
    readout shutter unit and calibration unit, with dedicated mechanical and thermal interfaces with PLM.
    The secondary mirror (M2) is mounted on a mechanism (M2M) for 3-DOF adjustment to compensate for launch and
    cool-down effects. In addition, the PLM hosts the FGS, used as pointing reference by the AOCS. All these detectors are
    mounted on the structure carrying the VIS focal plane, in order to ensure precise co-alignment.
    Except proximity electronics of the focal planes and FGS, all electronics are placed on the SVM to minimise thermal
    disturbances to the PLM.
 
 
 
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