SolO ISP Study

1
Integrated Science Payload for the Solar Orbiter
Mission Final Review

• ESTEC June 29th 2004

2
Study overview
3
Study challenges and main steps

• To reduce the mass budget by 25 in order to
  recover the payload mass assumption made for the
  system assessment study.
• Mass reassessment of instruments as described in
  PDD shows opposite conclusion!
• Clarification/homogeneisation/relaxation of
  resolution requirements
• 1 arcsec spatial resolution / 150 km pixel
  targetted for all high resolution instruments
• Allows to reduce instruments size from 1.5 m to 1
  m length
• Allows to come back within mass specification
• Allows to better deal with solar flux
• To deal with the SolO mission challenge of a
  complex suite of instruments for an ambitious
  journey toward the sun, at a cost in line with an
  ESA flexible mission.
• First system level iterations indicates that S/C
  for shortest cruise missions were too heavy
• Instrument size reduction
• allows to design more compact spacecrat
• Now mass compatible with shortest cruise mission,
  using SEP and direct Venus transfer
• Remote sensing and in-situ Instrument I/F
  clarifications/consolidation
• Allows to initiate system studies with
  consolidated data
• Allows to promote I/F standardisation, to pave
  the way for an efficient development

4
Study team organisation
Frederic FAYE
Frederic FAYE
Frederic FAYE
Christian STELTER
Omar EMAM
5
Study logic
6
Study schedule
ISP for
SolO
ISP for
SolO
24/09/2003
M5
M6
M1
M2
M3
M4
24/09/2003
M5
M6
M1
M2
M3
M4
PM1
PM2
FP
WM
KO
PM1
PM2
FP
WM
KO
MTR
MTR
WPA
Management Expertise
WPA
Management Expertise
WP1
Instrument Performance System Assessment
WP1
Instrument Performance System Assessment
110
Mission Spacecraft assessment
110
Mission Spacecraft assessment
120
Instruments performance system assessment
120
Instruments performance system assessment
130
Radiation EMC assessment
130
Radiation EMC assessment
WP2
Instrument Resource Reduction
WP2
Instrument Resource Reduction
210
Resource reduction synthesis
210
Resource reduction synthesis
220
Sensor architecture technologies
220
Sensor architecture technologies
230
Mechanical
-
thermal architecture technologies
230
Mechanical
-
thermal architecture technologies
240
Electrical
-
functional architecture technologies
240
Electrical
-
functional architecture technologies
250
ISP support
Bepi
heritage
250
ISP support
Bepi
heritage
WP3
Conceptual design of
Rrsource
efficient payload
WP3
Conceptual design of
Rrsource
efficient payload
310
ISP system engineering
310
ISP system engineering
320
Sensor architecture technologies
320
Sensor architecture technologies
330
Mechanical
-
thermal architecture technologies
330
Mechanical
-
thermal architecture technologies
340
Electrical
-
functional architecture technologies
340
Electrical
-
functional architecture technologies
WP4
Payload technology planning cost analysis
WP4
Payload technology planning cost analysis
WP5
Shared payload subsystems planning cost analysis
WP5
Shared payload subsystems planning cost analysis
7
Environment analyse

• Space enviromnent
• Contamination guidelines

8
Environment Analysis

• Source Term Mission Solar Proton Fluence

9
Environment Analysis

• Total Dose (Cruise Mission)

10
Environment Analysis

• Source Term Solar Wind
• Solar wind carry considerable kinetic energy,
  typically 1 keV for protons and 4 keV for He.
  This can result in sputtering from exposed
  surface materials
• Flux 1.3 E9 particles/(cm² s) (average),
  Momentum flux rv² very high gt1E16!!

11
Environment Analysis

• Radiation Effects and Consequences on SOLAR
  ORBITER P/L
• Degradation of electronic components, detectors
  due to ionising dose
• No significant problem for shielded (4mm)
  electronics and sensors (14 krad)
• Non ionising absorbed dose (displacement) due to
  protons
• Displacement in bipolar devices is an issue but
  generally negligible below about 3E10 p/cm² (50
  MeV)
• Displacement on optical devices (optocoupler,
  APS, etc.) very critical
• gt Solutions on parts level (hardening
  technology) and on system level (intelligent
  shielding is efficient),
• gt APS remain problematic
• Galactic Cosmic Ray induced effects (single event
  phenomena SEP)
• no further problem for SOLO compared to missions
  at 1AU w/o geomag. Shielding
• Solar event (proton and ion) induced upsets
  (single event phenomena SEP)
• A factor of 25 higher at 0.2 AU than in GEO
• Measures in order to cover the problem mainly on
  electronic design level (filtering, EDAC,
  TMR, etc.)

12
Environment Analysis

• Meteroid fluence on Solar Orbiter
• Design parameters v45 km/s, r2 g/cm³, impact
  angle 45

13
Environment Analysis

• Solar Dust exposure

14
Cleanliness Analysis

• EMC
• EMC Control requires normalEMC measures on S/C
  level
• EMC program/working group requested by RPW

15
Cleanliness Analysis

• Magnetic Cleanliness
• MAG requires magnetic cleanliness plan (TBD), but
  according to Science Teams response
  (Sci-A/2004/069/AO, 9/6/2004) no anticipated
  problems stated.

16
Cleanliness Analysis

• Particulate/Organic Cleanliness
• Cleanliness and Contamination Control follow
  ECSS-Q-70-01A
• Particulates
• Cleanroom conditions, e.g. CLASS 10 000 for PWA
  at all times
• Organic Cleanliness
• Materials not to be used
• polymeric materials with high outgassing
  potential
• polymeric materials with low particle radiation
  stability (radiolysis)
• Halogenated polymeric materials

17
Conclusions on environment and cleanliness

• Environment assessement
• Major care shall be taken against
• Displacement due to Proton (in particular with
  APS systems)
• Solar events (protons and ions) induced upset
• Solar wind effects (sputtering on thin layers)
• Material selection (radiolysis)
• No major concerns arise from total radiation dose
  and GCR
• Contamination assessement
• Cleanliness plan are needed for all payloads,
  covering
• EMC cleanliness
• Magnetic cleanliness
• Particulate organic cleanliness (outgassing)
• This will drive the allowable material list
• At system level, an evaluation of Suitability of
  an Integrated Shielding System (Thermal, MM Dust,
  Radiation) deserves consideration

18
Remote sensing instruments

• VIM

19
Visible-light Imager and Magnetograph
(VIM)Overview

• Measurement of
• velocity fields using Doppler effect
• magnetic fields using Zeeman effect
• Magnetograph imagery in narrow (5 pm FWHM )
  spectral bands around a visible spectral line at
  different polarisation states ? line of sight
  (LOS) velocity ? magnetic field vector
• Time resolution 1 minute (5 ? x 4
  polarisations)
• Spatial resolution
• 0.5 arc-sec with 0.25 arc-sec sampling ? 250 mm
  (PDD)
• 1 arc-sec with 0.5 arc-sec sampling ? 125 mm
  (new baseline)
• Field 2.7 (angular diameter of sun at 0.21 AU)
• Split in 2 instruments HRT for resolution and
  FDT for field
• Stringent LOS stability 0.02 arc-sec over 10 s
  (differential photometry) ? internal Image
  Stabilisation System

20
Visible-light Imager and Magnetograph
(VIM)Functional block diagram
HRT High Resolution Telescope
FO Filtergraph Optics
focus and image stabilisation mechanism
visible filter
PMP Polarisation Module Package
Fabry Perot in collimated beam
selection mirror
detector
aperture door mechanism
front end electronics
28 V
collimator
mechanism drive electronics
camera
back end electronics
limb sensor
FDT Full Disk Telescope
21
VIM configuration, volume and mass
resolution relaxation ? volume and mass reduction
PDD new design
HRT resolution sampling field diameter 0.5 arc-sec0.25 arc-sec8.5 arc-min250 mm 1 arc-sec0.5 arc-sec8.5 arc-min125 mm
FDT resolution sampling field diameter 9.5 arc-sec4.75 arc-sec2.726 mm 19 arc-sec9.5 arc-sec2.713 mm
focal planes 2k x 2k 1k x 1k
volume 1300 x 400 x 300 800 x 400 x 300
mass 30 kg (PDD)35.4 kg (Astrium) 30 kg (Astrium) (with 20 margin)
excluding window, enclosure radiators
22
Critical items and proposed alternatives

• Critical technologies and alternatives
• Polarisation Modulation Package 10-3
  polarisation accuracy, tuning?1s
• Liquid Crystal Variable Retarders behaviour
  under radiations
• alternative wheel mechanism with polarisers
• Fabry Perot FWHM 5 pm, FSR 150 pm, ?1s
• LiNbO3 solid state etalons with spectral tuning
  achieved by high voltages behaviour under
  radiations
• alternatives vacuum with piezo or thermal
  deformation, gaz with pressure control
• proposed demonstrators in technological plan
• Proposed VIM design modifications
• Narrow band entrance filter to minimize heat
• Off-axis optical configuration for HRT( to avoid
  strong obturation by heat stop)or refractive
  system

23
Remote sensing instruments

• EUS

24
EUV Imager and Spectrometer (EUS)Overview

• High resolution slit spectrometry of sun disk
• Three spectral bands
• 17 22 nm
• 58 63 nm
• 91.2 tbd nm
• Spatial resolution sampling 0.5 arc-sec (PDD)
  ? 1 arc-sec (new)
• Diameter 120 mm (? 60 mm) not driven by
  diffraction effects but by flux ? optics
  transmission is a key parameter (2 telescope
  options)
• Spectral resolution 1 pm/pixel (PDD) ? 2
  pm/pixel (new)
• Spectrometer concept single element toroidal
  varied line-space (TVLS) grating
• Field of view 34 arc-min driven by detector
  array size (4k ? 2k)
• Spectral range 4-5 nm driven by detector array
  size (4k ? 2k)
• Internal raster mode
• Internal LOS control system from VIM data (tbc)

25
EUV Imager and Spectrometer (EUS)Functional
block diagram
telescope single mirror or Wolter II
relay optics with disperser
proposed EUV filter
slit as field stop
detector
front end electronics
raster mode LOS control by mirror tilting
28 V
shutter
mechanism drive electronics
back end electronics
26
EUV Imager and Spectrometer (EUS)Recommandations

• Normal Incidence System (NIS) for the telescope
• EUS requires a large diameter entrance aperture
  (120 mm), leading to large solar heat loads,
  above 400 W at 0.21 AU ? Entrance EUV filter
  with radiative grid recommended
• Al foil filter well adapted for two bands

27
EUV Imager and Spectrometer (EUS)Radiative grid
on Al foil

• A radiative grille (black painted) parallel to
  Sun flux is conductively coupled to the metal
  filter, and allow to radiate the absorbed flux.
  The global emissivity of the filter assembly is
  highly increased.

28
EUS configuration, volume, mass
resolution relaxation ? volume and mass reduction
PDD new design
sampling field diameterspectral 0.5 arc-sec34 arc-min120 mm1 pm / pixel 1 arc-sec34 arc-min60 mm2 pm / pixel
focal plane 4k x 4k 2k x 2k
volume 1600 x 400 x 300 800 x 140 x 150
Mass(1) (2) 25 kg (PDD)31.8 kg (Astrium) 15.2 kg (with 20 margin)
(1) increase pixel to 8 µm would lead to a
volume of about 960 x 240 x 180 (2) ancillary
equipment, thermal cover not yet accounted for
29
EUS with relaxed resolutionThermal issue

• Proposed EUS design with relaxed resolution ? 60
  mm pupil diameter ? re-opening of entrance filter
  trade-off
• Option 1 pupil on mirror

30
EUS with relaxed resolutionThermal issue

• Option 2 pupil at instrument entrance
• Advantage reduced heat load on baffle
• Drawback oversized primary mirror, optical
  design to be reassessed

entrance diameter 60 mm 91.7 W
heat stop radiator59 to 67 Wto be rejected
31
EUV Imager and Spectrometer (EUS)Critical points
and open issues

• Option with entrance filter
• obturation of filter radiator impact on
  throughput
• EUV filter ? thermal issue is solved
• breadboard in technological plan
• Option without entrance filter (with reduced
  pupil)
• thermal control critical heat rejection of heat
  stop thermo-elastic deformationstypical
  tolerance 10µm / 100µrad ? 5C on SiC structure,
  some tenths of C on mirror gradients
• primary mirror multilayer coating behaviour with
  high thermal flux to be assessed
• EUV Detectors
• 2 k x 2 k format back-thinned CMOS with 5 µm
  (tbc) pixels
• breadboard in technological plan
• Toroidal varied-line gratings studies in US and
  Italy maturity of technology ?
• Coatings from 17 to 100 nm multilayer, gold, SiC
  2 or 3 bands ?

32
Remote sensing instruments

• EUI

33
EUV Imager (EUI)Overview

• Imaging of the sun disk in EUV
• Resolution/sampling 0.5 arc-sec (PDD) ? 1
  arc-sec (new)
• Field of view 2.7 (sun angular diameter at
  0.21 AU)
• Field/resolution 20 000 (?10 000) ? split in 2
  instruments
• HRI for resolution 0.5 arc-sec (? 1 arc-sec) in
  34 arc-min field (4k x 4k ? 2k x 2k detector
  array)
• FSI for field 4.75 arc-sec (? 9.5 arc-sec) in
  5.4 field (4k x 4k ? 2k x 2k detector array)
  field of FSI is twice the sun angular diameter to
  account for HRI depointing
• HRI spectral bands 13.3 nm, 17.4 nm, 30.4 nm ?
  3 different HRI telescopes optimised for each
  spectral band
• FSI spectral bands tbd in 17.1 30.4 nm ?
  single telescope
• Diameter of HRIs and FSI 20 mm driven by
  radiometry and not diffraction ? could be reduced
  to 10 mm with relaxed resolution
• Internal LOS control system from VIM data (tbc)

34
EUV Imager (EUI)Functional block diagram
EUV filter
telescope
relay optics
baffle
field stop
detector
front end electronics
aperture door mechanism
LOS control by mirror tilting
28 V
back end electronics
mechanism drive electronics
35
EUV Imager (EUI)Bafflage and EUV filter
HRI
FSI
36
EUV Imager (EUI)HRI and FSI configurations

• FSI
• baffle decoupled from optical bench
• filter supported by baffle

• HRI
• single structure ("optical bench") for all 3
  telescopes
• baffles thermally decoupled from the "optical
  bench" to minimise heat-flux and thermoelastic
  distortion

37
EUV Imager (EUI)Evolution of design
resolution relaxation ? volume and mass reduction
PDD new design
HRI sampling field diameter 0.5 arc-sec34 arc-min20 mm 1 arc-sec34 arc-min10 mm
FSI sampling field diameter 4.75 arc-sec5.420 mm 9.5 arc-sec5.410 mm
focal plane 4k x 4k 2k x 2k
volume 3 x 1800 x 450 x 150 1800 x 440 x 250 900 x 110 x 130 940 x 250 x 190
mass 42.6 kg (PDD)42.5 kg (Astrium) 14.6 kg
excluding window, enclosure radiators other
ancillary equipment
38
EUV Imager (EUI)Critical points and open issues

• Heat rejection of EUV filters and baffles
• EUV Detectors ( as EUS)
• back-thinned CMOS
• 4 k x 4 k ? 2 k x 2 k format with 9 µm pixels
• alternative detectors Diamond or GaN/AlGaN ?
  credible in large format ?
• Cooling of CMOS detectors at 80C
• Telemetry huge compression or data selection
  required

39
Remote sensing instruments

• COR

40
Coronograph (COR)Overview

• Observation of sun corona between 1.2 and 3.5
  radii
• Coronograph
• needs of occulters to mask the sun disk
• optical design with field stop and Lyot stop
• Spectral bands
• 450 600 nm
• 121.6 ? 10 nm
• 30.4 ? 5 nm (optional)
• Field of view 9.2 (corona angular diameter at
  0.21 AU)
• Spatial resolution spatial sampling 8 arc-sec
  driven by 4 k x 4 k detector array ? 16 arc-sec
  with 2 k x 2 k

41
Coronograph (COR)Functional block diagram
EUV/VIS dichroic
image internal occulter
achromatic polarimeter
UV filter wheel mechanism
external occulter
relay optics
entrance pupil
EUV detector
front end electronics
telescope
pupil Lyot stop
sun disk rejection mirror
aperture door mechanism
VIS detector
COR pointing mechanism
back end electronics
28 V
mechanism drive electronics
42
Coronograph (COR)Overall configuration
COR volume not in line with other remote sensing
instruments? recommandation decrease distance
external occulter to pupil with related decrease
of pupil diameter (at constant vignetting)
PDD new design()
sampling fielddiameter 8 arc-sec9.2 33 mm 16 arc-sec9.216.5 mm
focal plane 4k x 4k 2k x 2k
volume 1200 x 400 x 300 (PDD) 1400 x700 x 370 (Astrium) 800 x 400 x 250
mass 21.8 kg (PDD)40.7 kg (Astrium) 20 kg
To be assessed on science grounds
43
Coronograph (COR) Critical points and open issues

• Recommandations
• Mass and volume not in line with other remote
  sensing instrumentsrecommandation reduction of
  sampling distance ? shrinkage of instrument by
  factor 2
• removal of pointing mechanism ? COR off during
  offset pointing) ? duty cycle
• whole design to be worked out further
• Critical points and open issues
• Heat rejection of external occulter
• Design of sun disk rejection mirror
• EUV coating of mirrors compatible with visible
  light
• Feasibility of the EUV/VIS dichroic (visible
  light reflected, EUV get through)
• EUV detectors see EUS EUI

44
Remote sensing instruments

• STIX

45
Spectrometer Telescope for Imaging X-rays (STIX)
Overview

• Imagery of sun disk in X-rays
• Spectral bands hard X-rays 4 150 keV 8 -
  310 pm
• Use of X-rays techniques
• pseudo imaging by grids
• X-ray position/energy detectors CdZnTe
• Spatial resolution sampling 2.5 arc-sec
• Field of view
• FWHM imaging field of view 24 arc-min
• Spatially integrated spectroscopy field of view
  3
• Energy resolution 2 to 4 keV in 16 energy
  levels

PDD Reduction objective
volume 1500 x 70 x 70 1000 x 70 x 70
mass 5 kg 5 kg
46
Spectrometer Telescope for Imaging X-rays (STIX)
Functional block diagram
filter
Fresnel lens
aspect system
VIS detector
X-ray detector
front end electronics
rear grid
aspect system
front grid
28 V
back end electronics
47
Spectrometer Telescope for Imaging X-rays (STIX)
Critical points and open issues

• Good level of maturity in general
• Becomes the longest remote sensing instrument
  following reduction exercise on all other
  instruments
• Length reduction down to 1 m should be
  investigated, in line with dimensions of all
  other remote sensing instruments.
• This may require to reduce the grid pitch if
  resolution needs to be kept
• Will avoid to constrain the S/C design snowball
  impact on structure mass at S/C structure level,
  on orbiter and on propulsion module
• Aspect system design to be investigated further
• Detector CdZnTe space qualified prototype but
  design to be adapted to STIX

48
Synthesis for remote sensing instruments
49
Why integration of payloads is not practical for
Solar Orbiter ?

• Common telescope with shared focal plane ex
  Hubble, JWST, Herschell
• astronomy of faint objects very high resolution
  ? large pupil
• reduced spectral range Hubble visible, JWST
  IR, Hershell submillimeter
• small instruments with respect to collector

Solar orbiter reduced collector diameter (sun
at 0.2 AU) large instrument dimensions from
visible to X-rays very specific instruments
coronograph no possibility and no interest to
share optics ? conclusion no suite ? only
individal instruments
50
Remote sensing instruments geometrical IRD

• From the mechanical configuration, IRD are
  updated
• Overall volume including bipodes
• Electronic not included. Sizing based on DE
  boards
• Volume for connectors, closure box not included

PDD PDD PDD Updated Updated Updated
Length Width Height Length Width Height
VIM 1300 400 300 800 400 300
EUS 1600 400 300 800 140 150
3 x HRI 1800 450 150 900 110 130
FSI 1800 440 250 940 250 190
COR 1200 400 300 800 400 250
STIX 1500 70 70 1000 70 70
51
Remote sensing instruments mass budget

• Hypotheses for remote sensing intruments mass
  estimate
• Filter outside instruments not included
• Enclosures for protection against
  pollution/contamination not included
• Aperture doors for LEOP and may be SEP phases
• Instrument internal covers or enclosures for AIT,
  LEOP and outgassing phases
• Electronic masses not challenged
• Note Ancillary equipment / instrument enclosures
  not yet accounted for

initial PDD design initial PDD design design with relaxed resolution
PDD Astrium estimate design with relaxed resolution
STIX 5.0 5.0 5.0
VIM 30.0 35.4 30.0
EUS 25.0 31.8 15.2
EUI 42.6 42.5 14.6
COR 21.8 40.7 20.1
Total 124.4 155.4 107.5
52
Remote sensing instruments data rate
raw data and processing needs allocation
VIM Raw data rate Frame 1k x 1k, 3.2 s, 12 bits 3.75 Mbit/sAfter processing by FPGA computation of 5 parametersmax 1k x 1k, 3 parameters, 300 s, 4 bits ? 40 kbit/s peak 1k x 1k, 3 parameters, 60 s, 4 bits ? 200 kbit/s 20 kbit/s
EUS Raw data frame 2k x 2k, 1.27s frame, 12 bits 38 Mbit/s After selection and processing 6 lines, 3 line profile parameters, length 1000 pixels, 1.27 s frame, 12 bits, compression 1/10 17 kbit/s 17 kbit/s
EUI HRI raw data 3 HRI, frame 2kx2k, 10 s, 12 bits 14.4 Mbit/sWavelet compression with factor 48 ? 300 kbit/s (baseline)compression / selection scheme to TBD FSI raw data frame 2k x 2k, 4800 s, 12 bits 10 kbit/s 20 kbit/s
COR Raw data frame 2k x 2k, occultation 0.5, 600 s, 16 bits, factor 1.5 (UV1, vis0.5) 80 kbit/sLossless compression with factor 3 26.7 kbit/sadditional lossy compression of 5 required 5 kbit/s
STIX raw data 64 pixels, 16 energy channels, 8 Hz, 16 bits 130 kbit/s ? 1 hour storage in a 64 Mbyte rotating buffer after processing total count on 8 bits 64 relative values on 4 bits 56 bits miscellaneous 320 bits/imagex 1800 images/h (6 mn flare, 0.5 Hz, 10 energies) 25 other data 720 kbit/hour 0.2 kbit/s 0.2 kbit/s
53
Remote sensing instruments thermal aspects
PDD end of ISP study
VIM 2 options with and without window heat stop in optical beam narrow band filtering entrance window M1 radiator with radiative coupling suggested off axis optical configuration for heat stop heat stop radiator M2 screenwith radiative coupling optical bench thermal control
EUS heat stop at primay focal planepossibility of adaptive optics Al foil with radiative grid proposed at instrument entrance Open instrument back in the picture thanks to reduced aperture surface
EUI long baffles with vanes EUV filter unchanged
COR sun-disk rejection mirror thermal washers to decouple external occulter from structure low emissivity conical shapes on external occulter radiator coupled to sun-disk rejection mirror by fluid loop
STIX opaque sunshade 1 mm of carbon or 3 mm of Beryllium thin reflective coating on grids unchanged
54
Remote sensing instruments review outcomes

• All instruments appear feasible
• Alternative solutions have been identified for
  all identified critical items
• Potential science impact of alternative solutions
  to be assessed by science teams
• Limiting the thermal flux inside instrument was a
  driver for our assessement Open issue limited to
  EUS entrance filter, for which TDA are deemed
  mandatory impact on science
• PDD mass estimates are rather optimistic and not
  exhaustive (ancillary equipment), resolution
  relaxation (to be accepted by science tyeam) is
  proposed
• to reduce volumes and masses
• As a side effect, to limit the solar heat flow to
  deal with
• No show stopper for the mission, however payload
  mass/volume plays a critical role in the context
  of Solar Orbiter Assessment, as larger mass
  allocations imply longer cruise or mission
  profiles not in line with ESA flexi budget

55
Solar Orbiter Remote Sensing instruments
Technology plan

• EUV detector
• design and realisation of the basic technological
  elements for a large array , small pixel
  operating in visible and EUV
• performance and environment tests
• Radiative grid for EUV filter
• trade-off on material
• manufacturing and integration of EUV filter and
  grid
• thermal test
• Polarisation modulation package
• trade-off on technologies
• design of the package
• breadboard manufacturing
• environment tests
• Fabry Perot package
• trade-off on technologies
• design of the package
• breadboard manufacturing
• environment tests

56
Solar Orbiter visible and EUV focal planes
not blind EUV detectors
visible detectors
blind EUV detectors
MCP
vis Si
EUV Si
GaN/diamond
vis Si
CMOS
CMOS
CMOS
CMOS
or
or
or
vis APS monolithic
MCP
EUV APS monolithic
vis APS monolithic
C3PO
CMOS
standard CMOS
existing
RT ESA hybrid 18 µm
technology to be promoted
planned RT ESA
all detectors of Solar Orbiter require CMOS
CMOS
57
Solar Orbiter visible and EUV focal planesToday
status

• Visible detector
• hybrid CMOS as baseline to optimise quantum
  efficiency x fill factor
• monolithic CMOS as back-up
• C3PO requested ?
• Not blind EUV detector
• hybrid CMOS as baseline to make EUV optimisation
  easier
• EUV monolithic CMOS as back-up
• Blind EUV detector
• GaN hybridised on CMOS read-out circuit as
  baseline if technological development successful
• Otherwise MCP with visible detector
• Assumptions for technology plan
• GaN / diamond ESA RT is confirmed
• C3PO ESA RT is confirmed

58
Solar Orbiter visible and EUV focal
planesTechnology approach

• Statement for Solar Orbiter
• CCD not suitable because of irradiations
• CMOS is mandatory for all detectors of Solar
  Orbiter
• format 2k x 2k or 1k x 1k with 10µm pitch
• CMOS development must be secured and commonalised
  (cost reduction)
• selection of one CMOS technology (design rule,
  founders, CIS if monolithic) according to
  performances and irradiations hardening
• evaluation and qualification of this CMOS
  technology for Solar Orbiter
• develop guidelines for design of CMOS function
  with respect to irradiation hardening
• Transfer ESA RT  hybrid CMOS  from 18 to 10 µm
  pitch ? breadboard
• Optimisation of hybrid CMOS technology from
  visible to EUV ? breadboard
• In parallel, development of EUV monolithic APS
  (RAL development)

59
EUV filter with radiative grid

• Phase 1 9 months
• trade-off on material for grid major
  criterionmanufacturing polishing
  capability(optical surface for thin foil
  contact ? SiC good candidate
• manufacturing of the radiative grid
• assembly with foil (procurement)
• Phase 2 12 months with 3 months overlap
• thermal test on solar vacuum facility
• challenge simulate 25 solar constants? afocal
  telescope to be developpedwith cooling of
  secondary mirror
• cold space simulated by shrouds
• temperature of thinn foilmonitored with infrared
  camera
• test objectives check foil temperature
  correlate thermal model
• facility can be used to test other Solar Orbiter
  units

60
Polarisation Modulation and Fabry Perot packages

• Polarisation modulation package 18 to 24 months
• trade-off on technologies (tests on Lyquid
  Crystal to Solar Orbiter environment already
  performed)
• design of the package
• 2 retarders linear polariser
• oven with active thermal control
• breadboard manufacturing
• fonctionnal, optical and environment tests
• Fabry Perot package 18 to 24 months
• trade-off on technologies (tests on Lithium
  Niobate to Solar Orbiter environment already
  performed)
• design of the package
• 2 Fabry Perots 1 interference filter
• oven with active thermal control
• breadboard manufacturing
• fonctionnal, optical and environment tests

61
In situ instruments

• Plasma package

62
SWA

• Composed of three types of sensors
• Electron Analyser Sensor (EAS),
• Proton Alpha particle Sensor (PAS)
• Heavy Ions Sensor (HIS).
• They are characterised by
• Their large field of view requirements,
• EAS Electrons coming from every directions
• PAS HIS Particle incidence driven by magnetic
  field
• The need to operate below 40C
• PAS and HIS have to Sun pointed
• Accommodated directly behind the Sunshield
• With a collector in direct Sun light
• Main issues
• EAS accommodation on P/F, boom or body mounted
• HIS and PAS collector in Sunlight
• Should be decoupled from rest of instrument,
• Should be coupled to S/C structure for thermal
  control
• Should reasonably not exceed 10 cm2 (i.e. 30 W
  load per head)

SWEA on Stereo (EAS)
SWICS on Ulysses (HIS)
Triplet on Interball (PAS)
63
SWA EAS

• 2 heads body mounted provides a quasi 4 p Sr
  coverage

64
SWA HIS and PAS in SunlightHigh conductance
device candidates

• Several type of light high conductance devices
  are possible candidates for coupling heat loaded
  zone to cold radiators

1. Mini fluid loop
Total mass 80 g Global conductance 1 W/K for
up to 30 W Distance Heat source / radiator up
to 50 cm Flight tested in 2003 and 2004
(COM2PLEX, Ariane5 ECA in summer 2004)
Evaporator (?25 mm x 19 mm)
Condensor (mounted on a radiator)
2. Conductive strap
Graphite fiber thermal straps
Copper or aluminium straps
65
In Situ instruments

• Field package
• RPW
• CRS
• MAG

66
RPW

• Interface accommodation requirements mainly
  characterised by
• The three 5 m long electric antennas, to be
  accommodated possibly in Sunlight and orthogonal
  to each others,
• What is the material considered for the antennas?
  
• The loop magnetometers and the search coil
  magnetometers, to be accommodated away from the
  spacecraft on deployable boom,
• The need to operate magnetometer sensors below
  50C, i.e. protected from the direct Sun flux,
• A clean EMC environment although not yet
  quantified for operations.

67
CRS

• Makes use of the spacecraft communication system,
  
• X band uplink
• Dual band X / Ka downlink
• Possibly complemented by a lightweight Ultra
  Stable Oscillator
• The physical accommodation constraints will then
  be limited to define a proper compromise for the
  USO location between
• minimum harness length, thus close to TRSP
• clean and stable environment (thermal, EMC), thus
  far from TRSP.
• Main issues
• Found a suitable location inside location for USO
• Define USO thermal control stability requirement
• The reference mission profile does not provide
  actual solar conjunctions
• Open question radio science compliance with
  simultaneous TM downlink?

68
CRSSun spacecraft Earth angle over the
mission
Launch Cruise Nominal mission Extended
mission
69
MAG

• Interface accommodation requirements
  characterised by
• The need to implement the sensors away from the
  spacecraft body on long deployable boom,
• The need to maintain the sensors below 57C, i.e.
  protected from the direct Sun flux,
• A clean EMC environment although not yet
  quantified for operations,
• Rosetta approach -characterisation only- seems
  not sufficient
• Cluster approach is too demanding and not
  compatible with Bepi euse
• The need to slew the spacecraft at several deg/s
  about the Sun direction to regularly calibrate
  the fluxgate magnetometer.

70
In Situ instruments

• Particle package
• EPD
• DUD
• NGD

71
EPD

• Includes five sets of sensors
• Supra-Thermal Electron detector (STE),
• Electron and Proton Telescope (EPT),
• Supra-thermal Ion Spectrograph (SIS),
• Low Energy Telescope (LET)
• High Energy Telescope (HET).
• Interface accommodation requirements
• The large FOV requirements, requiring either
• rotating platform
• multiple sensor option,
• The need to operate below 30C, i.e protected
  from the direct Sun flux,
• Main issue
• FOV blockage by S/C body in case of scan P/F

72
DUD

• Interface accommodation requirements
  characterised by
• the need to hard mount two small units on the
  spacecraft side
• and provide them with wide /- 80 free FOV
• One unit to be mounted in the orbital plane 90
  off the S/C-Sun line
• The other perpendicular to the orbit plane

73
NGD

• Interfaces accommodation requirements
  characterised by
• Sun pointed instrument, below a shield window no
  thicker than 3g/cm2
• Sensors kept below 30C, i.e. protected from the
  direct Sun flux.

74
In Situ instruments

• Main issues

75
In situ instruments Main issues

• Instrument design
• Low resources demands,excepted FOV
• Sensors rather well defined
• Sharing of electonics widely suggested
• Instruments accommodation
• All sensors but RPW electrical antennas and SWA
  PAS/HIS collectors to be placed behind sun shield
• Instrument environments
• Most instruments deemed EMC sensitive, but no
  cleanliness specification (apart RPW sensitivity)
  on the table to date
•  Good design practices  claimed to be
  sufficient

76
Instrument accommodation

• Trade off overview

77
Expected effects of resources reduction options


Resource
reduction option

Communalisation of
Technology
Standardisation

Development
Resource

functions

improvements

centralisation



Mass

à

à

à
?

à
à
à


Power




consumption


Volume

à

à


à


Data storage and
à








data rate

à
Þ
à
à
Development cost






Devel
opment time

Þ


Þ

à



78
Architecture options trade off

• Mechanical-thermal design

79
Opto mechanical alternatives

• No clear advantages of integrated design
• Preferred solution depends on relative weighting
  between mass and integration
• Considering the major configuration differences
  between instruments
• Individual instrument design kept as baseline

80
Accommodation at spacecraft level
Accommodation on
spacecraft
Central
cylinder
Polygonal
shape
Plane
assembly
Pros
Cons
Pros
Cons
Cons
Pros

Stiffness

Launch
axis

Stiffness

Launch
axis

Stiffness

Flexible

Alignment
constraint

Alignment
constraint

Alignment
geometry

Weight

Structure
inter planes
driver
Payload module favoured
Integrated design

• Mechanical accommodation to be addressed at
  spacecraft level
• No reason to impose a payload module in the frame
  of ISP study

81
Trade offs for steerable platform

• Need for steerable platform
• At ISP level, platform is the preferred solution
• Compatibility with pointing stability requiremnts
  to be confirmed at spacecraft level.

82
Platform geometry options and trade off
83
General thermal control principle

• Remote sensing instruments
• Instruments are thermally controlled
  independently
• A maximum of Sun flux is stopped at the entrance
  of each instruments thanks to
• customised filters, (EUV, visible),
• heat stops (at intermediate focus point),
• radiating baffles,
• rejection mirror (for coronograph),
• Once the totality of the main part of Sun flux is
  stopped, telescopes and optical bench require
  classical thermal control (several lines of
  heaters, thermistors, ON/OFF or PID law). They
  benefit of radiators shadowed by the Sunshield
  (very stable environment).
• Detectors are independently thermally controlled.
  A dedicated radiator with a good thermal coupling
  (flexible strap or fluid loop) is foreseen
• Location of instruments and their radiators
  (detectors, telescope, heat loaded areas) is to
  be coupled with satellite configuration study
  (possible view factor with solar arrays and/or
  back side if the sunshield).
• In situ instruments
• When possible protected by thermal shield
• Minimise surfaces in direct Sunlight
• Ecouple Sun illumintaed surface from the rest of
  the instrument

84
Remote sensing instuments PDD status for filters
85
Remote sensing instruments Thermal issue
recommendation

• Recommendation is to try systematically to
  minimise unneeded heat load inside instruments
• Implement filters
• Outside instruments,
• Coupled to S/C wall or sunshield
• Develop large EUV filters, in particular because
  EUV bandwidth is marginal wrt heat flux
• Thermal control becomes no more a critical
• For instruments
• For instruments/SC interfaces
• To be dealt with in dedicated instrument
  assessement studies
• This statement is reinforced with the smaller
  apertures resulting from the revised resolution

86
Recommended filter implementation summary
(1) depending whether an adequate filter
material can be found for EUS
87
Instruments cover

• Covers needed
• To avoid contamination deposits and risk of
  polymerisation under UV flux
• Launch, LEOP, propulsion phases
• To avoid solar flux entry during slight
  offpointing (COR)

88
Need and possible location for Sun pointed
instruments covers
89
WP 200 Architecture options trade off

• Pointing pointing stability for remote sensing
  P/L

90
Pointing constraints

• Pointing direction
• EUI/FSI, EUS, VIM/HRT and STIX need spacecraft
  off pointing to cover Sun disk
• VIM FDT off when VIM HRT operates
• EUI/FSI  oversized  to cope with off pointing
• COR needs
• Option 1 mechanism
• Option 2 switch off and cover during off
  pointing
• Pointing stability
• VIM requires a very high pointing stability
• EUI/EUS call for 0.1 arcsec/s class performance
• Not achievable using standard S/C systems
• Option 1 Instruments image stabilisation system
• Close loop system as VIM complex but OK
• Open loop system (EUI/EUS) questionable (S/C
  behaviour)
• gt Option 2 Post processing on ground

91
Alternatives to meet pointing stability
requirements
Pointing
stabilisation options
Pointing
stabilisation options
4
4
1
2
3
5
1
2
3
5
6
7
6
7
Two
instruments
Two
instruments
One instrument (VIM)
One instrument (VIM)
Each
remote sensing
Each
remote sensing
One instrument (VIM)
Spacecraft
Spacecraft
(VIM STIX)
One instrument (VIM)
Spacecraft
Spacecraft
(VIM STIX)
provides
provides
Instrument
detects
Instrument
detects
provides
provides
guarantees
provides
provides
provides
guarantees
provides
error stability
signal
error stability
signal
Image
reconstructed
Image
reconstructed
pointing errors and
pointing errors and
high pointing stability
high pointing stability
error stability
signal
pointing error
signal
error stability
signal
error stability
signal
pointing error
signal
error stability
signal
to
other
instruments
to
other
instruments
on
ground
on
ground
compensates with
compensates with
as
required
by
as
required
by
to
spacecraft
AOCS
as
required
by
to
other
instruments
to
spacecraft
AOCS
as
required
by
to
other
instruments
equipped with
equipped with
After
post
processing
After
post
processing
its own
image
its own
image
all instruments
all instruments
which controls
all instruments
equipped with
which controls
all instruments
equipped with
their own
image
their own
image
Apart VIM provided with its own closed loop
stabilisation system
stabilisation system
stabilisation system
attitude
accordingly
excepted
VIM
including
VIM
their own
image
attitude
accordingly
excepted
VIM
including
VIM
their own
image
stabilisation system
stabilisation system
stabilisation system
stabilisation system

Spacecraft with

Spacecraft with

Spacecraft with

Spacecraft with

Spacecraft with

Spacecraft with

Spacecraft with

Spacecraft with
low
performance
low
performance
high
performance
high
performance

Spacecraft with
high
performance
low
performance

Spacecraft with
high
performance
low
performance

Spacecraft with

Spacecraft with

Spacecraft with

Spacecraft with
AOCS
and
AOCS
and
AOCS
with
AOCS
and
AOCS
and
AOCS
with
AOCS
and
AOCS
and
low
performance
low
performance
very high
performance
high
performance
very high
performance
high
performance
inter P/L DMS
inter P/L DMS
simple DMS
simple DMS
AOCS
and
P/L in
the loop
interP
/L DMS
AOCS
and
P/L in
the loop
interP
/L DMS
AOCS
and
AOCS
and
AOCS
and
AOCS
and
simple DMS
simple DMS
simple DMS
simple DMS
simple DMS
simple DMS

Only
VIM
with

Only
VIM
with

Only
VIM
with

Only
VIM
with

Only
VIM
with

Two
instruments

Only
VIM
with

Two
instruments
with error detection
with error detection
with error detection
with error detection

Complex
instruments
with error detection
with error detection

Complex
instruments
with error detection
with error detection

Simple instruments

Simple instruments

Simple instruments

Simple instruments
Image stabilisation
Image stabilisation
Image stabilisation
Image stabilisation
with error detection
Image stabilisation
Image stabilisation
with error detection
Image stabilisation
Image stabilisation
without pointing
without pointing
without pointing
without pointing
systems

other
systems

other
systems

other
systems

other
Image stabilisation
systems

other
systems

other
Image stabilisation
systems

other
systems

other
systems
systems
systems
Systems (but VIM)
Instruments
without
Instruments
without
Instruments
without
Instruments
without
Instruments
without
Instruments
without
Instruments
without
Instruments
without
systems
systems
Image stabilisation
Image stabilisation
mechanism
mechanism
mechanism
mechanism
mechanism
mechanism
systems only
systems only
92
WP 200 Architecture options trade off

• Instrument data management

93
Partitionning drivers

• Very high raw data rate
• Limited use of dedicated Gbps point-to-point
  links
• First stage of data reduction in the instrument
  front-end
• Instrument-specific high performance computation
• Hardwired implementation, not shareable, possibly
  expandable for size reasons
• Compression
• Discrepancy between raw data volume achievable
  and downlink capability requires to clarify
  compression schemes, duty cycles, or even
  instrument concepts
• Standard implementation of a (multiple) data flow
  processing chain
• Flexible algorithms
• Communalised algorithms to be find out
• Thermal Control
• Instrument led fine thermal control of inner
  hardware parts
• Best at front-end level for AIV reasons
• Specific processing
• Case by case analysis
• High degree of flexibility, at least during the
  implementation phase

94
Partitionning baseline(remote sensing
instruments)
95
Partitionning baseline(in-situ instruments and
augmentation)
96
Architectural design driversPayload interface

• Guideline
• One interface per instrument electronic module
  (FEE or MDE),
• Merging of science and control command data
• No SPF impacting more than one instrument
• High rate links concentrators close to
  instruments to reduce harness
• Science interface
• Standard Spacewire links (even if one way high
  rate only required)
• Bepi Colombo solutions promoted
• Control command interface
• Based on SpW micro-remote terminal unit derived
  from ESA TDA

97
Traded solutions and recommended one
HICDS with P/F I/Os SpW-only PDPU µRTU for
P/L
HICDS with P/F I/Os bi-bus PDPU µRTU for P/L
HICDS with I/Os bi-bus PDPU
HICDS with I/Os monobus PDPU
Fully centralized
HICDS PDPU generalised µRTU sensor bus
HICDS with P/F I/Os PDPU with CAN
HICDS with P/F I/Os multi-purpose PDPU
HICDS with P/F I/Os Generic PDPU P/L RTU
HICDS PDPU RTU for all I/Os
98
WP 200 Architecture options trade off

• Instruments power distribution

99
Instrument power needs
100
Alternative power design solutions
Power distribution alternatives
Distribution of regulated primary power
Centralised distribution of Primary and secondary
power
Distributed standard converters
101
Traded power distribution solutions
Power bus from PDU to switchable instruments
Power bus from PDPU to switchable instruments
Individual protected lines from PDU to switchable
instrument
Individual protected lines from PDU to non
switchable instruments
Individual protected lines from CPPS to non
switchable instruments
Individual protected lines from PDU to switchable
instruments grouped per suite and location
Individual protected lines from PDPU to
switchable instrument
Individual protected lines from PDPU to non
switchable instruments
102
Instrument accommodation

• Recommended baseline

103
Possible accommodation for in situ payloads
104
Baseline electrical interface
Data management
Power distribution
Standard
converter
Standard
converter
28 V
regulated
28 V
regulated
PCDU
PCDU
LCL
CV
LCL
CV
LCL
LCL
On/Off command
On/Off command
P/L or P/L suite
Space Wire
Space Wire
I/O,
I/O,
PDPU
PDPU
TM/TC
TM/TC
Micro RTU
Micro RTU
Standard electrical interface
105
The scan platform for EPD sensors

• Characteristics
• Mass about 2 kg
• Power about 1 W
• 2 Mrpm over 6 years
• Encoder 1 deg accuracy
• 2 DE boards electronics
• Development
• 1 QLTM 1 PFM
• 28 months incl 6 months FB

106
Overall payload mass budget
107
Instrument assessment

• The Unionics assessment

108
WHY UNIONICS?

• Modular Design
• Can utilise a common modular design approach
  replicated across nodes.
• Semi-mass production reducing cost and
  schedule.
• One type of module one type of test set-up.
• Seemless transfer of functions across nodes
  without having to shutdown.
• Simple High Speed Interconnect
• High speed Space wire currently 200Mbit/s
  projected 3.5Gbit/s.
• A number or off the shelf Space Wire routers
  are now available.
• Can form redundant connections and easily
  isolate faults.
• Can by pass faulty nodes and re-route data and
  commands.
• Well established protocols and routing software.
• DSP21020 (software option)
• DSP MCM mature space qualified design used on
  INM4
• Current MCM can operate at speeds of 14MHz
  achieving 20MIPs
• Built in space wire interfaces.
• Mature software for multi-tasking across a
  network of DSPs.
• Software able to reconfigure network of DSPs and
  redistribute run time programs as necessary.
• FPGA (hardware option)
• Large (1Mgate) very high speed space qualified
  FPGAs are now available.

109
How Can UNIONICS Apply to SOP?

• SOP Multiple instruments
• Potentially similar front end electronics
  interfaces.
• Distributed system with instruments spread over
  the platform.
• SOP Very high throughput of raw data
• Requiring data processing and compression,
  closely related to the front end node.
• Need for high speed links between node (including
  Mass Memory).
• SOP Require accurate pointing information
• Observation and attitude control are closely
  inter-related.
• Can use UNIONICS concept to extend the payload
  data processing and. overlap with attitude
  monitoring and control by including them as
  additional nodes.
• SOP Requiring Mass Memory (MM)
• Space qualified MM of up to 800Gbits are being
  manufactured by ASTRIUM.
• MM access speed of up to 400Mbit/s can be
  achieved.
• Interface to MM can be easily adapted to be
  compatible with Space Wire
• MM can effectively appear as a UNIONICS node in
  the system.

110
Assumptions for SOP payload DPU (Cont.)

• Centralised Box
• Advantages
• Can be specified prior to completion of payload
  instruments definition.
• Independent of payload instruments design,
  manufacture and test.
• Oversized generic design which could be reused on
  other platforms.
• Internal modular design so that DPU can be down
  sized if necessary.
• Standard multiple but duplicate SpW I/Fs.
• Integrated power conditioning (reduced packaging
  and power losses).
• Integrated mass memory (reduced packaging and
  interconnect requirements).
• Disadvantages / Problems
• Harness and connectors mass may be large.
• Harness routing may be problematic.
• Accommodating a large mass and volume unit on a
  small platform.
• Maintaining low temperatures for a small unit
  volume dissipating high power.
• Existing mass memory module mechanical design may
  have to be modified.
• Existing power conditioning module mechanical
  design may have to be modified.

111
Assumptions for SOP payload DPU

• Space Wire (SpW) Interconnect
• Advantages
• Industry standard interface, approved and
  supported by ESA.
• SpW is an inherently reliable and fault tolerant
  interconnect architecture.
• Already base lined for SOP and Bepi-Columbo.
• Availability of Space qualified ASICs, IPs and
  other building blocks.
• Availability of generic routing software.
• Availability of UNIONICS software which utilises
  SpW.
• Availability of of the shelf prototyping and test
  equipment and software.
• Good EMC performance.
• High data throughput, gt200Mbit/s.
• Disadvantages / Problems
• Four-core differential interconnect, higher g/m
  than some other alternatives.
• Not as efficient in terms of Bitrate/MHz/W as
  some alternative dedicated links.
• Continuos token and clock required on active
  interfaces.
• Point-to-point interconnect requiring the
  overhead of
• Switching Matrix(s)

112
Assumptions for SOP payload DPU (Cont.)

• Centralised 300Gbits SDRAM Mass Memory
• Advantages
• Based on 100Gbit (2 x 50Gbits independent banks)
  Module, a DC/DC converter and a Chip Set for Mass
  Memory control with high fault coverage.
• Already manufactured for Pleiades.
• File management capability.
• No software required for SSR control.
• Design can withstand cosmic radiation dose of up
  to a 100Krad.
• High latchup LET threshold (TBC awaiting final
  test).
• Low SEU susceptibility (TBC awaiting final
  test).
• At the maximum scrubbing rate a LEO SEU rate as
  low as 10-17 error/bit/24h (TBC).
• Small volume per module (13x250x250 mm