integrated science payload for the solar orbiter mission final review

SolO ISP Study – FR - ESTEC – 29 June 2004 SOP-HO-ASF-023

Integrated Science Payload for the Solar Orbiter Mission Final ReviewESTEC– June 29th 2004


Study overview


Study challenges and main steps

1. 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

2. 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


Study team organisation

Contractual officer

Christine DURAND - A s triu m S A S

E lectrical - functiona larchitectu re & tec hnologie s

L u c PL ANCHE - Astrium SAS

ISP support & Bepi heritag e

M arc STECKL ING - Astrium G m bH

Radiations & EM C assessm en t

W o lfg ang KEIL - Astrium G m bH

Unionics a ssessm en t

An d y CRO SS - Astrium Ltd

M ission & Integrated SciencePayload (ISP) engineering

Eric M ALIET - A s triu m S A S

In-s itu instrum ents consultanc y

Jean -An d ré SAUVAUD - CESR

Independen t Scientific consultan tSolar an d helios pheric m ission s

Pr Rain er SCHW EN N

Rem ote-sensing inst rum ents consultanc y

Peter R . YO UNG - RA L

Instrum entsassessm ents

Dom inique DUBET - A s triu m S A S

Payload Technology & Subsytemsplanning and cost analysis

Eric M ALIET - A s triu m S A S

SolO Integrated Science PayloadsStudy Manager

Eric M ALIET - A s triu m S A SFrederic FAYE

Frederic FAYE Frederic FAYE

Christian STELTER

Omar EMAM


Study logic

Kick Off

WP 1

Instrument Performance & System Assessment

Mission & spacecraft assessment

WP 3

Conceptual design of Resource Efficient IPS

WP 4

Technology development plan & cost analysis

WP 2

Instrument resource reduction options

Trades-off, and identification of preferred approach

WP 5

Shared P/L subsystemsdevelopment plan& cost analysis

PM 1

PM 2

M T R

Final Review

Kick Off

WP 1



WP 3


WP 4


WP 2



WP 5


WP 1



WP 3


WP 4


WP 2



WP 5

Shared payloadsubsystemsdevelopment plan& cost analysis

PM 1

PM 2

M T R

Final Review

Kick Off

WP 1



WP 3


WP 4


WP 2



WP 5


WP 1



WP 3


WP 4


WP 2



WP 5


PM 1

PM 2

M T R

Final Review

Kick Off

WP 1



WP 3


WP 4


WP 2



WP 5


WP 1



WP 3


WP 4


WP 2



WP 5


PM 1

PM 2

M T R

Final Review

WP 1



WP 3


WP 4


WP 2



WP 5


WP 1



WP 3


WP 4


WP 2



WP 5


PM 1

PM 2

M T R

Final Review


Study schedule

ISP forSolO

WPA Management & Expertise

WP1 Instrument Performance & System Assessment110 Mission & Spacecraft assessment120 Instruments performance & system assessment130 Radiation & EMC assessment

WP2 Instrument Resource Reduction210 Resource reduction synthesis220 Sensor architecture & technologies230 Mechanical-thermal architecture & technologies240 Electrical-functional architecture & technologies250 ISP support &Bepi heritage

WP3 Conceptual design ofRrsource efficient payload310 ISP system engineering320 Sensor architecture & technologies330 Mechanical-thermal architecture & technologies340 Electrical-functional architecture & technologies

WP4 Payload technology planning & cost analysisWP5 Shared payload subsystems planning & cost analysis

24/09/2003 M5 M6M1 M2 M3 M4

PM1 PM2 FPWMKO MTR

ISP forSolO

WPA Management & Expertise

WP1 Instrument Performance & System Assessment110 Mission & Spacecraft assessment120 Instruments performance & system assessment130 Radiation & EMC assessment

WP2 Instrument Resource Reduction210 Resource reduction synthesis220 Sensor architecture & technologies230 Mechanical-thermal architecture & technologies240 Electrical-functional architecture & technologies250 ISP support &Bepi heritage

WP3 Conceptual design ofRrsource efficient payload310 ISP system engineering320 Sensor architecture & technologies330 Mechanical-thermal architecture & technologies340 Electrical-functional architecture & technologies

WP4 Payload technology planning & cost analysisWP5 Shared payload subsystems planning & cost analysis

24/09/2003 M5 M6M1 M2 M3 M4

PM1 PM2 FPWMKO MTR


Environment analyseSpace enviromnentContamination guidelines


Environment Analysis

Source Term: Mission Solar Proton Fluence



Total Dose (Cruise + Mission)



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 E+9 particles/(cm² s) (average), Momentum flux ~ v² very high >1E+16!!



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 3E+10 p/cm² (50 MeV)– Displacement on optical devices (optocoupler, APS, etc.) very critical

=> Solutions on parts level (hardening technology) and on system level (intelligent shielding is efficient), => 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.)

Interference with detector operation (background produced by secondary nuclear reactions)– Thorough analysis on proton interaction with materials (surface material, shielding structure) and evaluation of activation effects

(spallation, neutron, gamma emission)

Radiation induced outgassing (radiolysis) and following contamination– Selection of non polymeric with non-halogenic content materials

Solar Wind Effects– Evaluation of solar wind degradation effects (sputtering) on surface materials (change of , surface roughness, etc.)



Meteroid fluence on Solar Orbiter– Design parameters: v=45 km/s, =2 g/cm³, impact angle 45°

SOLAR ORBITER Meteoroid Fluence9y extended mission duration, =2g/cm³

1,00E-07

1,00E-06

1,00E-05

1,00E-04

1,00E-03

1,00E-02

1,00E-01

1,00E+00

1,00E+01

1,00E+02

1,00E+03

1,00E+04

1,00E+05

1,00E+06

1,00E-06 1,00E-05 1,00E-04 1,00E-03 1,00E-02 1,00E-01 1,00E+00

Diameter [cm]

Flu

en

ce

[#

/m²]



Solar Dust exposure


Cleanliness Analysis

EMC EMC Control requires

„normal“EMC measures on S/C level

EMC program/working group requested by RPW

INSTRUMENT/UNIT EMC REQUIREMENT

REMARK Ref.

Plasma Package PDD Iss. 2 SWA (EAS, PAS, HIS)

Normal S/C, not sensitive

Field Package PDD Iss. 2 RPW S/C wide EMC

program, as outlined in the separate EMC doc

EMC doc requested!

CRS N/A MAG No special

requirement for EMC but for magnetic cleanliness

Magnetic cleanliness plan (TBD)

Particle Package PDD Iss. 2 EPD Not specified DUD Not specified NGD None special Remote Sensing Instruments

PDD Iss. 2

VIM TBD EUS Not stringend Normal S/C

equipment levels

EUI Not stringend Normal S/C equipment levels

COR Not specified STIX None special High Priority Augmentation

PDD Iss. 2

DPD Not specified NPD Surface of

detector (TBC)

RAD N/A



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.

INSTRUMENT/UNIT MAGNETIC REQUIREMENT

REMARK Ref.


not specified

Field Package PDD Iss. 2 RPW Not specified CRS N/A MAG < 1 nT Magnetic

cleanliness plan (TBD)

Particle Package PDD Iss. 2 EPD Not specified DUD TBD NGD No strong B-field

near operating unit

B-field could affect PMT ( field strength TBD)

Remote Sensing Instruments

PDD Iss. 2

VIM TBD EUS Not stringend Normal S/C

equipment levels

EUI Not stringend Normal S/C equipment levels

COR Not specified STIX None special High Priority Augmentation

PDD Iss. 2

DPD TBD NPD Not required RAD N/A



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

INSTRUMENT/UNIT PARTICULATE REQUIREMENT

MOLECULAR REQUIREMENT

REMARK


Class 10 000 cleanroom at all times

MCPs very sensitive (water, hydrocarbon, etc.)

Purging through testing and in fairing highly desirable

Field Package PDD Iss. 2 RPW Not specified Not specified Electric antennas

can produce some particulates during deployment

CRS N/A N/A MAG No special req. No special req. Particle Package PDD Iss. 2 EPD Not specified, Dry

N2 purging during ground operation

Not specified, Avoid acids, organic liquids (exception ethanol),

DUD Not specified, Dry N2 purging during ground operation

Not specified

NGD None special None special Remote Sensing Instr.

PDD Iss. 2

VIM TBD TBD Open vs filter has impact

EUS BOL 85 ppm EOL 150 ppm

BOL 5.0E-8 g/cm² EOL 1.0E-7 g/cm²

SOHO levels required

EUI BOL 85 ppm EOL 150 ppm

BOL 5.0E-8 g/cm²- Level: A/20 EOL 1.0E-7 g/cm²

SOHO levels required

COR Dust free, avoid light scattering from optics

Organic free avoid photo-polymerisation

STIX None special None special High Prior. Augment.

PDD Iss. 2

DPD Not specified Not specified NPD Class 10 000

cleanroom (TBC) Not specified

RAD N/A N/A Cleanliness as for other instruments


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


Remote sensing instrumentsVIM


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


Visible-light Imager and Magnetograph (VIM)Functional block diagram

detector

front endelectronics

back endelectronics

HRT: High Resolution Telescope

FDT: Full Disk Telescope

FO: Filtergraph Optics

Fabry Perot in collimated beam

28 V

aperture door mechanism

visible filter

PMP : PolarisationModule Package

collimator

camera

selectionmirror

limb sensor

mechanism drive electronics

focus andimage stabilisation

mechanism


VIM configuration, volume and mass

PDD new design

HRT resolution sampling

field diameter

0.5 arc-sec0.25 arc-sec8.5 arc-min

250 mm

1 arc-sec0.5 arc-sec8.5 arc-min

125 mm

FDT resolution sampling

field diameter

9.5 arc-sec4.75 arc-sec

2.7°26 mm

19 arc-sec9.5 arc-sec

2.7°13 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

resolution relaxation volume and mass reduction


Critical items and proposed alternatives

Critical technologies and alternatives– Polarisation Modulation Package : 10-3 polarisation accuracy, tuning1s

· 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

0

50

100

150

200

250

300

350

-100 -50 0 50 100 150 200

Zerodur

Silicacte = 5 10-6

Le

Zerodur

Silicacte = 5 10-6

Le


Remote sensing instrumentsEUS


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)


EUV Imager and Spectrometer (EUS)Functional block diagram

shutter

detector


back endelectronics

28 V

proposedEUV filter

telescopesingle mirroror Wolter II

slit asfield stop

relay opticswith disperser

raster mode & LOS control by mirror tilting



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


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.

4.5 mm

0.5 mm Alu foil 0.3

micron

Alu radiator

240 mm sunshie

ld

Satellite structure

VDA for absorption limit

EUV is transmitted

Visible and UV are mainly reflected

High coupling with cold space

0.3 mm substrat


EUS configuration, volume, mass

PDD new design

sampling

field

diameter

spectral

0.5 arc-sec34 arc-min

120 mm1 pm / pixel

1 arc-sec34 arc-min

60 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


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

700 mm

pupil on mirror diameter 60 mm

entrancediameter 67 mm

114 W

33.6 W on baffle

80.4 W on mirror

mirror radiator8 to 16 W

to be rejected

10% to 20%absorbed

80% to 90%reflected

59 to 67 W absorbed by heat stop

heat stop radiator59 to 67 W

to be rejected5 W inside spectro


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

entrancediameter 60 mm

91.7 W

heat stop radiator59 to 67 W

to be rejected

700 mm

mirror diameter 67 mm

11.3 W on baffle

80.4 W on mirror

mirror radiator8 to 16 W

to be rejected

10% to 20%absorbed

80% to 90%reflected

59 to 67 W absorbed by heat stop

5 W inside spectro

pupil at entrance diameter 60 mm


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 deformations

typical tolerance 10µm / 100µrad 5°C 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 ?


Remote sensing instrumentsEUI


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)


EUV Imager (EUI)Functional block diagram

aperturedoor

mechanism

detector


back endelectronics

28 V

EUV filtertelescope

field stop

relay optics

LOS control by mirror tilting


baffle


EUV Imager (EUI)Bafflage and EUV filter

HRI

FSI

entrance pupildiameter = 20 mm

foil filterdiameter = 35 mm

10 W

7 W absorbed directly by baffle2.7 W absorbed after reflection on foil

3 W reach foil filter0.3 W absorbed by

foil2.7 W reflected back

towards baffle

1500 mm

0.3° field

entrance pupildiameter = 20 mm

foil filterdiameter = 133 mm

10 W10 W reach the foil

filterspread on 66 mm

1 W absorbed by foil9 W reflected back

towards baffle 1190 mm

2.7° field

9 W absorbed by baffleafter reflection on foil filter


EUV Imager (EUI)HRI and FSI configurations

HRI :– single structure ("optical bench")

for all 3 telescopes – baffles thermally decoupled from

the "optical bench" to minimise heat-flux and thermoelastic distortion

FSI :– baffle decoupled from

optical bench– filter supported by baffle


EUV Imager (EUI)Evolution of design

PDD new design

HRI sampling field

diameter

0.5 arc-sec34 arc-min

20 mm

1 arc-sec34 arc-min

10 mm

FSI sampling field

diameter

4.75 arc-sec5.4°

20 mm

9.5 arc-sec5.4°

10 mm


volume3 x 1800 x 450 x

1501800 x 440 x 250

900 x 110 x 130940 x 250 x 190

mass* 42.6 kg (PDD)42.5 kg (Astrium) 14.6 kg

* excluding window, enclosure & radiators& other ancillary equipment



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 – 80°C

Telemetry: huge compression or data selection required


Remote sensing instrumentsCOR


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


Coronograph (COR)Functional block diagram

EUV/VISdichroic

EUV detector


back endelectronics

VIS

detecto

r

external occulter entrance

pupil

telescope

imageinternal

occulter

relayoptics

pupilLyot stop

UV filterwheel

mechanism

sun disk rejection

mirror

aperturedoor

mechanism


COR pointingmechanism

28 V

achromaticpolarimeter


Coronograph (COR)Overall configuration

PDD new design(*)

sampling

fielddiamet

er

8 arc-sec9.2 °

33 mm

16 arc-sec9.2°

16.5 mm


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

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)

* To be assessed on science grounds


Coronograph (COR) Critical points and open issues

Recommandations:– Mass and volume not in line with other remote sensing instruments

recommandation : 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


Remote sensing instrumentsSTIX


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


Spectrometer Telescope for Imaging X-rays (STIX) Functional block diagram

X-ray detector


back endelectronics

28 V

VIS detector

aspect system

aspect systemfront grid

rear grid

Fresnel lensfilter


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


Synthesis for remote sensing instruments


Why integration of payloads is not practical for Solar Orbiter ?

Common telescope with shared focal plane : ex Hubble, JWST, Herschellastronomy of faint objects + very high resolution large pupilreduced spectral range ; Hubble = visible, JWST = IR, Hershell =

submillimetersmall instruments with respect to collector

inst 1

inst 3

inst 2

inst 4

flux collector

Solar orbiter : reduced collector diameter (sun at 0.2 AU)

large instrument dimensions ; from visible to X-raysvery specific instruments: coronograph

no possibility and no interest to share optics conclusion : no suite only individal instruments


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 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


Remote sensing instruments mass budget

initial PDD designdesign with

relaxedresolutionPDD Astrium

estimate

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

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


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 parameters

max: 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/sAfter 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 TBDFSI 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 (UV=1, vis=0.5) = 80 kbit/s

Lossless compression with factor 3: 26.7 kbit/s;additional 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


Remote sensing instruments thermal aspects

PDD end of ISP study

VIM2 options : with and without windowheat stop in optical beam

narrow band filtering entrance windowM1 radiator with radiative couplingsuggested off axis optical configuration for heat stop + heat stop radiatorM2 screenwith radiative couplingoptical bench thermal control

EUSheat stop at primay focal planepossibility of adaptive optics

Al foil with radiative grid proposed at instrument entranceOpen instrument back in the picture thanks to reduced aperture surface

EUI long baffles with vanes + EUV filter unchanged

COR

sun-disk rejection mirrorthermal washers to decouple external occulter from structure

low emissivity conical shapes on external occulterradiator coupled to sun-disk rejection mirror by fluid loop

STIX

opaque sunshade: 1 mm of carbon or 3 mm of Berylliumthin reflective coating on grids

unchanged


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


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


Solar Orbiter visible and EUV focal planes

vis Si

CMOS

EUV Si

CMOS

GaN/diamond

CMOS

vis Si

CMOS

MCP

vis APSmonolithic

EUV APSmonolithic

visibledetectors

not blind EUVdetectors

blind EUV detectors

vis APSmonolithic

MCP

or or or

CMOSC3PO

standard CMOS

planned R&T ESA

R&T ESAhybrid 18 µm

CMOS

existingtechnologyto be promoted

all detectors of Solar Orbiter require CMOS


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 R&T is confirmed– C3PO ESA R&T is confirmed


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 R&T « 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)


EUV filter with radiative grid

Phase 1: 9 months– trade-off on material for grid: major criterion:

manufacturing+ 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 foil

monitored with infrared camera– test objectives: check foil temperature

+ correlate thermal model– facility can be used to test other Solar Orbiter units

4.5 mm

0.5 mmAlu foil 0.3

micron

Alu radiator

240 mm

sunshield

Satellite structure


EUV is transmitted



0.3 mm substrat

4.5 mm

0.5 mmAlu foil 0.3

micron

Alu radiator

240 mm

sunshield

Satellite structure


EUV is transmitted



0.3 mm substrat

solarconstant

1m

25 solarconstant 0.2m


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


In situ instrumentsPlasma package


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 40°C 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)

Triplet on Interball (PAS)

SWICS on Ulysses (HIS)


SWA EAS

2 heads body mounted provides a quasi 4 Sr coverage

2 EAS sensors are sufficient to coverfull space in absence

of obstacle

nearly4 PI FOV providedwith2 EAS sensors

on a boom away from S/C

SC body

S/C induces FOVblocking for sensorson

scanning platform Need to combine

with S/C yaw manoeuvres

or witha single one when combinedwith S/C yaw manoeuvres

S/C body

2 body mounted EAS sensors required to achieve 4 pi FOV

S/C body

or a single one when combinedwith S/C yaw manoeuvres

EAS sensoraccommodation

Body mounted Platform mounted Boom mounted

S/C bodyS/C body


of obstacle


of obstacle


of obstacle

nearly4 PI FOV providedwith2 EAS sensors

on a boom away from S/C

SC body

S/C induces FOVblocking for sensorson

scanning platform Need to combine

with S/C yaw manoeuvres


S/C body


S/C bodyS/C body

2 body mounted EAS sensors required to achieve 4 pi FOV

S/C bodyS/C body

or a single one when combinedwith S/C yaw manoeuvres

EAS sensoraccommodation

Body mounted Platform mounted Boom mounted

S/C bodyS/C bodyS/C bodyS/C body


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 :

Evaporator (25 mm x 19 mm)

Condensor (mounted on a radiator)

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)

2. Conductive strap

Graphite fiber thermal strapsCopper or aluminium straps


In Situ instrumentsField packageRPWCRSMAG


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 50°C, i.e. protected from the direct Sun flux,

– A clean EMC environment although not yet quantified for operations.


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?


CRSSun – spacecraft – Earth angle over the mission

0153045607590

105120135150165180

0 500 1000 1500 2000 2500

Days from launch

Sun

-Spa

cecr

aft-S

un a

ngle

(deg

)

Launch :Cruise :Nominal mission :Extended mission :


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 57°C, 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.


In Situ instrumentsParticle packageEPDDUDNGD


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 30°C, i.e protected from the direct Sun flux,

Main issue– FOV blockage by S/C body in case of scan P/F


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


NGD

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


In Situ instruments Main issues


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


Instrument accommodationTrade off overview


Expected effects of resources reduction options

Resource reduction option

Resource Communalisation of

functions Technology

improvements Standardisation Development

centralisation

Mass = ?

Power consumption

=

Volume =

Data storage and data rate

= = =

Development cost

Development time =


Architecture options & trade offMechanical-thermal design


Opto mechanical alternativesRemote sensing instruments

Opto-mechanical accommodation options

Instruments sharing same

primary structure

Combined instruments Separated instruments


Telescope

Instruments with common footprint IF

Instruments without peculiar

constraints

1 2 3 4 5

Remote sensing instrumentsOpto-mechanical accommodation options


primary structure

Combined instruments Separated instruments


Telescope

Instruments with common footprint IF

Instruments without peculiar

constraints

1 2 3 4 5

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


Accommodation at spacecraft level

Accommodation on spacecraft

Polygonal shape Central cylinder Plane assembly

Pro’s•Stiffness•Alignment

Con’s•Launch axis

constraint•Weight

Pro’s•Stiffness•Alignment

Con’s•Launch axis

constraint•Structure driver

Pro’s•Flexible

geometry

Con’s•Stiffness•Alignmentinter planes

Sun

dire

ctio

n

Sun

dire

ctio

n

Laun

ch d

irect

ion

Laun

ch d

irect

ions

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

Payload module favoured Integrated design


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.

Options PROs CONs

Rotating platform

- Minimisation of resources (mass : -6kg, power : -7W, data rates, cost ?)

- improved measurements angular resolution

- favors an integrated payload approach

- Limited FOV clearance (2/3 of required FOV)

- poor time resolution for 3D distributions

- burdensome to other instruments (pointing stability, electromagnetic noise)

Multiple body-fixed

sensor heads

- best time resolution - severe increase of required resources (mass, power, data rates)

- might raise S/C accommodation problems to satisfy all FOVs


Platform geometry options and trade off

Criteria Weig

ht

Long

cyl

ind

er

Short

cyl

ind

er

U-s

hap

ed

pla

tform

No p

latf

orm

Accessible FOV 1 1 0 -1 2Instantaneous FOV 1 1 1 0 2Mass 3 1 1 0 -2Harness 1 2 2 2 -2Data rate 1 0 0 0 -2P/F complexity 1 -1 -1 0 1S/C I/F complexity 2 -2 -1 -1 0

Total mark 2 3 -1 -5

Scanning platform for HSIS suite

Long cylinder Short cylinder U-shape mounting No ptlatform

Scanning platform for HSIS suite

Long cylinder Short cylinder U-shape mounting No ptlatform


General thermal control principle

Remote sensing instruments1. Instruments are thermally controlled independently

2. 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),

3. 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).

4. Detectors are independently thermally controlled. A dedicated radiator with a good thermal coupling (flexible strap or fluid loop) is foreseen

5. 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

1. When possible protected by thermal shield

2. Minimise surfaces in direct Sunlight

3. Ecouple Sun illumintaed surface from the rest of the instrument


Remote sensing instuments: PDD status for filters

Th

erm

al sh

ield

S/C

str

uct

ure

Inst

rum

en

t en

tran

ce

Befo

re o

pti

cs

Aft

er

pri

mary

m

irro

r

1 2 3 4 5 6

VIM (*) X X (*) Preferred optionEUI HRI X After baffle and vanes => limited flux

FSI X After baffle and vanes => limited fluxEUS XCOR X XSTIX X

No filter described

Remark

Outside instrumentInside instument

Filter position

On thermal shield On S/C structure At instrument entrance Inside instrument

Before optics After primary mirror

No filter

1 2 3 4 5

Filter position

On thermal shield On S/C structure At instrument entrance Inside instrument

Before optics After primary mirror

No filter

1 2 3 4 5


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


Recommended filter implementation summary

The

rmal

sh

ield

S/C

str

uctu

re

Inst

rum

ent

entr

ance

Bef

ore

opti

cs

Aft

er

prim

ary

mir

ror

1 2 3 4 5 6

VIM X (X)Further heat stop between primary and secondary mirror in an off axis configuration

EUI HRI XBaffles, vanes and Al foil thermally decoupled from optical bench

FSI XBaffles, vanes and Al foil thermally decoupled from optical bench

EUS X (1) X(1)

Trade off between stand alone radiative grid and conductive grid coupled to spacecrfat radiators to be carried out at spacecraft level.

COR X

The external occulter and the Sun rejection disk remains attached to the instrument which is mounted on a mechanism to compensate spacecrfat off pointing required by narrow filed of view instruments

STIX X

As for VIM, it is recommended to decouple the entrance filter from the instrument and to couple it to the spaccrfat structure

Remark

Outside instrument Inside instument No filter

(1): depending whether an adequate filter material can be found for EUS


Instruments cover

Cover location

Mounted in front of the instrument Inside instrument

Mounted on S/C structure

Mounted on Sun shield

1 2 3 4

Cover location

Mounted in front of the instrument Inside instrument

Mounted on S/C structure

Mounted on Sun shield

1 2 3 4

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)

Cover location

Mounted on the instrument

Mounted on S/C structure or

Sun shield

Tiltable RotableTiltable Rotable

Collective Individual

Cover location

Mounted on the instrument

Mounted on S/C structure or

Sun shield

Tiltable RotableTiltable Rotable

Collective Individual


Need and possible location for Sun pointed instruments covers

Possible cover position (from figure /

)

Instrument Cover needed

1 2 3 4

Remark

VIM-HRT Y X - X X Cover needed to protect entrance glass filter

VIM-FDT Y X - X X Cover needed to protect entrance glass filter

EUS Y/N X X X X Y : Needed in case of open telescope N: Not needed with entrance thermal filter

EUI-HRI TBC X X X X TBC with Al foil filter on tip of entrance baffle

EUI-FSI TBC X X X X TBC with Al foil filter on tip of entrance baffle

COR Y X X X X Needed for the body mounted COR during off pointing Sun observation sessions

STIX N X - - - Protection provided by the front grid

SWA-PAS Y X - X X Cover needed to protect detector

SWA-HIS Y X - X X Cover needed to protect detector

NGD N - - - X Protection provided by the Sun shield itself


WP 200: Architecture options & trade offPointing & pointing stability for remote sensing P/L


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)

=> Option 2: Post processing on ground To be investigated


Alternatives to meet pointing stability requirements Pointing stabilisation options

Each remote sensing Instrument detects pointing errors andcompensates with its own image

stabilisation system

One instrument (VIM)provides

error stability signalto other instruments

equipped with their own image



error stability signal to spacecraft AOCS

which controls attitude accordingly

Spacecraft provides

pointing error signalas required by all instruments excepted VIM

Spacecraft guarantees

high pointing stabilityas required by all instruments including VIM

1 2 3 4 5

• Spacecraft withlow performance AOCS andsimple DMS

• Complex instrumentswith error detection& Image stabilisation

systems

• Spacecraft withlow performance AOCS andinter P/L DMS

•Only VIM withwith error detection& Image stabilisation

systems; other Instruments without Image stabilisation systems only

• Spacecraft withhigh performance AOCS with P/L in the loop


systems; other Instruments without mechanism

• Spacecraft withhigh performance AOCS andsimple DMS

• Only VIM withwith error detection& Image stabilisation


• Spacecraft withvery high performance AOCS andsimple DMS

• Simple instrumentswithout pointing

systems

6

Two instruments (VIM + STIX)

provides error stability signal

to other instrumentsequipped with

their own image stabilisation system

• Spacecraft withlow performance AOCS and

interP /L DMS

• Two instrumentswith error detection& Image stabilisation


Image reconstructed on ground

After post processing



systems

7

Pointing stabilisation options

Each remote sensing Instrument detects pointing errors andcompensates with its own image



error stability signalto other instruments

equipped with their own image



error stability signal to spacecraft AOCS

which controls attitude accordingly

Spacecraft provides

pointing error signalas required by all instruments excepted VIM

Spacecraft guarantees

high pointing stabilityas required by all instruments including VIM

1 2 3 4 5

• Spacecraft withlow performance AOCS andsimple DMS

• Complex instrumentswith error detection& Image stabilisation

systems

• Spacecraft withlow performance AOCS andinter P/L DMS


systems; other Instruments without Image stabilisation systems only

• Spacecraft withhigh performance AOCS with P/L in the loop




• Only VIM withwith error detection& Image stabilisation


• Spacecraft withvery high performance AOCS andsimple DMS


systems

6

Two instruments (VIM + STIX)

provides error stability signal

to other instrumentsequipped with

their own image stabilisation system

• Spacecraft withlow performance AOCS and

interP /L DMS

• Two instrumentswith error detection& Image stabilisation


Image reconstructed on ground

After post processing



Systems (but VIM)

7

Apart VIM provided with its own closed

loop stabilisation system


WP 200: Architecture options & trade offInstrument data management


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


Partitionning baseline(remote sensing instruments)

P/L Front-end data processing FEE output (PDD) FEE output (reduced resolution)

Back-end data processing (PDPU)

VIM Computation of 5 physical magnitudes using FPGA Computation of the pointing error at a frequency 100

Hz. Distribution to EUI and EUS may require a dedicated link. Synchronisation signal provided by spacecraft. ICU gathering local control, including instrument thermal

control

50 kbps (mode 1,3) 100 kbps (mode 2) 500 kbps (mode 4) 2.4 Mbps (peak)

12.5 kbps (mode 1,3) 25 kbps (mode 2)

125 kbps (mode 4) 0.6 Mbps (peak)

Lossless compression Lossy compression Very high compression rate (wrt

scientific return)

EUS Selection Line parameters computing (need CPU ?)

Image differencing to be studied (raw data rate: 670 Mbps - 10mn cadence, 2000 steps, 200Mb/image)

Synchronisation signal provided by spacecraft. ICU gathering local control, including instrument thermal

control

170 kbps

42.5 kbps

Lossy compression

EUI HRI : None (baseline) – Acquisition freq may increase, leading to an ideal raw data rate of 3*560 Mbps (1680 Mbps) requiring local processing like e.g. selection through image processing, image differencing (still to be studied) or even wavelet compression (1:48).

FSI : None (baseline, one image required every 4800s) – Potential image processing for HRI data selection

Synchronisation signal provided by spacecraft. ICU gathering local control, including instrument thermal

control

3 * 4.8 Mbps (HRI baseline)

4.8 Mbps

(FSI peak from PDD)

3 * 1.2 Mbps (HRI baseline)

4.8 Mbps

(FSI peak from PDD)

Lossy compression HRI: Very high compression

rate (wrt scientific return)

COR None Assuming 50% of images loaded every 600s/1200s (full

image: 270 Mb)

220 kbps +

110 kbps

No change Lossless compression Lossy compression

STIX Accumulation (2048 bits every 1/8 sec) 128 kbps No change Specific processing: processing

power need not critical


Partitionning baseline(in-situ instruments and augmentation)

Instrument Front-end data processing FEE output Back-end data processing (DPU) SWA-EAS None 2 Mbps burst Lossy compression

Data reduction SWA-PAS None 800 kbps burst Lossy compression

Data reduction SWA-HIS FPGA based integration algos (heritage TBC) 200 kbps + 500

bps Lossy compression Data reduction

EPD Hardwired processing (accumulation, sample selection, etc.) for all sensors STE, EPT, SIS, LET, HETn

256 kbps, from 5 sensors

Specific processing: Data reduction

DUD Hardwired processing (TBC) 50 bps x2 None NGD FPGA based data reduction (classification) (heritage TBC)

Mechanical collocation 400 bps

up to 15kbps Lossy compression TBC

RPW Waveform dectection thru FFT & correlations for TNR and RAD TBD processing for the 10 Mbps LFR data ICU gathering all data processing with direct output to the SSMM.

Few kbps None

MAG None (raw data) Mechanical collocation in HBIS electronics box

6.5 kbps Data/events provided to other instruments (TBC and low frequency) Specific processing: low processing need assumed

Augmentation instruments DPD Specific reduction TBD TBC TBD Lossless compression NPD None TBD Low input rate assumed

Specific processing: low processing need assumed RAD heritage TBC 3.6 kbps Specific processing: processing power need not critical


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


Traded solutions and recommended one

HICDS

PayloadArea

SpW router

I/O

TFG

SpW router

SSMM

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s (

MIL

-ST

D-1

55

3B

)

P/L Processor

Platform

HICDS

PayloadArea

SpW router

I/O

TFG

SpW router

SSMM

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s (

MIL

-ST

D-1

55

3B

)

P/L Processor

Platform

HICDS

PayloadArea

I/O

TFG

SSMM

PDPU

SpW router

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

Platform

SpW routerHi rate links (SpW)

HICDS

PayloadArea

I/O

TFG

SSMM

PDPU

SpW router

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

Platform

SpW routerHi rate links (SpW)

HICDS

PayloadArea Spw router

I/O

TFG

SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

Platform

Low rate science bus (CAN)

Can IF

HICDS

PayloadArea Spw router

I/O

TFG

SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

Platform


Can IF

HICDS

PayloadArea

Spw router

TFG


SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

RTU

RTU

RTU

RTU

I/O

Platform

Can IF

HICDS

PayloadArea

Spw router

TFG


SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

RTU

RTU

RTU

RTU

I/O

Platform

Can IF

RTU

HICDS

PayloadArea

Spw router

I/O

TFG

SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

Platform


Can IF

RTU

HICDS

PayloadArea

Spw router

I/O

TFG

SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

Platform


Can IF

Sys

tem

bu

s

P/L RTU

PayloadArea

Spw routerSpW router

SSMM

PDPU

Hi rate links (SpW)

HICDS TFG

Bus IF

I/O

SpW router

CAN gateway


I/OPlatform

Sys

tem

bu

s

P/L RTU

PayloadArea

Spw routerSpW router

SSMM

PDPU

Hi rate links (SpW)

HICDS TFG

Bus IF

I/O

SpW router

CAN gateway


I/OPlatform

PDPU

HICDS

PayloadArea

I/O

TFG

Low rate bus (CAN)

SpW router

SSMM

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

I/O

Platform

Can IF

PDPU

HICDS

PayloadArea

I/O

TFG

Low rate bus (CAN)

SpW router

SSMM

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

I/O

Platform

Can IF

Lo

w r

ate

sc

ien

ce

&

co

ma

nd

/co

ntr

ol

bu

s (

CA

N)

PDPU

HICDS

PayloadArea

TFG

SpW router

SSMM

Hi rate links (SpW)

Can IF

RTU

RTU

RTU

RTU

RTU

I/O

Bus IF

Platform

Lo

w r

ate

sc

ien

ce

&

co

ma

nd

/co

ntr

ol

bu

s (

CA

N)

PDPU

HICDS

PayloadArea

TFG

SpW router

SSMM

Hi rate links (SpW)

Can IF

RTU

RTU

RTU

RTU

RTU

SensorBus I/F

Platform

RTU

RTU

…

…

Bus I/F

I/O

Lo

w r

ate

sc

ien

ce

&

co

ma

nd

/co

ntr

ol

bu

s (

CA

N)

PDPU

HICDS

PayloadArea

TFG

SpW router

SSMM

Hi rate links (SpW)

Can IF

RTU

RTU

RTU

RTU

RTU

SensorBus I/F

Platform

RTU

RTU

…

…

Bus I/F

I/O

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

Fully centralized HICDS with I/Os + monobus PDPU

HICDS with I/Os+ bi-bus PDPU

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

RTU

HICDS

PayloadArea

Spw router

TFG

SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

RTU

RTU

I/O

Platform

SpW router

RTU

RTU

RTU Hi rate links (SpW)

PacketWire (SpW)

(TRSP)

UIBSpaceWire

RTU

HICDS

PayloadArea

Spw router

TFG

SpW router

SSMM

PDPU

Hi rate links (SpW)

Bus IF

Sys

tem

bu

s

RTU

RTU

I/O

Platform

SpW router

RTU

RTU

RTU Hi rate links (SpW)

PacketWire (SpW)

(TRSP)

UIBSpaceWire


WP 200: Architecture options & trade offInstruments power distribution


Instrument power needs

Instrument Power

Need (W) Specific power supply Mechanisms

In-Situ Instruments

SWA 11 30 kV from +2.4 kV to +3.9 kV from +8.6 V to +300 V from -2.4 kV to -3.9 kV from -15 V to -2 kV

Yes

RPW 6.4 Deployment

CRS 3 No

MAG 1.5 No

EPD 8.3 ~4 kV Yes

NGD 4 0-2.0 kV, 1.3 kV typical No

DUD (x2) 1 No

DPD 3 from 100 V to 4 kV Yes

NPD 2 TBD

Remote Sensing Instruments

VIM 34 2 kV 20 V low frequency

Yes

EUS 25 Yes

EUI 20 No

COR 34 High voltage Yes

STIX 4 No

RAD 7.4 Yes


Alternative power design solutions

Power distribution alternatives

Distribution of regulated primary power

Payload Distribution Unit

LCL

LCL

LCL

Instrument X power Supply

Unit

DC/ DC CV

Primary Power bus (28V) Secondary Power supplies

Central Payload Power Supply

LCL

LCL

LCL

DC/ DC CV

Primary Power bus (28V)

Secondary Power supplies

V1

Vn

To Inst. 1

To Inst. x

Instrument X power Supply Unit

LCL DC/ DC

CV

Primary Power bus (28V) Secondary Power supplies

Standard Converter

Centralised distribution ofPrimary and secondary power Distributed standard converters


Traded power distribution solutions

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV

HOIS + STIXStandard CV

PDPUStandard CV

PDU

28 V DC power bus

Standard CVVIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

PDU

28 V DC power bus

Standard CV

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


SC PDU

PDPU

28 V DC power bus

Standard CVVIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


SC PDU

PDPU

28 V DC power bus

Standard CV VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

Standard CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

Standard CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSDC/DC CV

EUIDC/DC CV

CORDC/DC CV

HBISDC/DC CV

HSISDC/DC CV

HSPISDC/DC CV

HOIS + STIXDC/DC CV

PDPUDC/DC CV

DC/DC CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSDC/DC CV

EUIDC/DC CV

CORDC/DC CV

HBISDC/DC CV

HSISDC/DC CV

HSPISDC/DC CV

HOIS + STIXDC/DC CV

PDPUDC/DC CV

DC/DC CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


S/C PDU

Standard CV

PDPU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


S/C PDU

Standard CV

PDPU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSDC/DC CV

EUIDC/DC CV

CORDC/DC CV

HBISDC/DC CV

HSISDC/DC CV

HSPISDC/DC CV

HOIS + STIXDC/DC CV

S/C PDU

DC/DC CV

PDPU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSDC/DC CV

EUIDC/DC CV

CORDC/DC CV

HBISDC/DC CV

HSISDC/DC CV

HSPISDC/DC CV

HOIS + STIXDC/DC CV

S/C PDU

DC/DC CV

PDPU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

PDU

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

Standard CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

PDU

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

Standard CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

CPPS

VIM

EUS

EUI

COR

HBIS

HSIS

HSPIS

HOIS + STIX

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

LLC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

DC/DC CV

HVCV

HVCV

HVCV

HVCV

HVCV

CPPS

VIM

EUS

EUI

COR

HBIS

HSIS

HSPIS

HOIS + STIX

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

LLC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

DC/DC CV

HVCV

HVCV

HVCV

HVCV

HVCV

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 PDPU to switchable instrument

Individual protected lines from PDPU to non switchable instruments

Individual protected lines from PDU to switchable instruments grouped per suite and location

Individual protected lines from CPPS to non switchable instruments


Instrument accommodationRecommended baseline


Possible accommodation for in situ payloads

HBIS suite•MAG•RPW mags•SWA EAS

HSIS (pos 2)•EPD

HSPIS•SWA PAS, HIS•NGP

RPW antenna

DUD (DPD)(NPD)

VelocitySun

Normal to orbit plane

HSIS (pos 1)•EPD HBIS suite

•MAG•RPW mags•SWA EAS

HSIS (pos 2)•EPD

HSPIS•SWA PAS, HIS•NGP

RPW antenna

DUD (DPD)(NPD)

VelocitySun


VelocitySun


HSIS (pos 1)•EPD


Baseline electrical interface

EUI MDE

EUI ICU

EUS MDE

EUS ICU

EUS Electronics Box

EUI Electronics Box

VIM MDE

VIM ICU

VIM Electronics Box SpW switch

DPU Core(LEON)

Payload Data Processing Unit

SpW router

SSMM

Routing Unit

SWA-HIS FEE

SWA-PAS FEE

HSPIS Electronics Box

NGD FEE

RTU

STIX FEE

DUD FEE

HOIS/STIX Electronics Box

SWA-EAS FEE

RTU

COR MDE

COR FEE

COR Electronics Box

RTU

RTU

RTU

RTU

HSIS Electronics Box

P/F common

SIS FEE

HETn FEE

LET FEE

EPT FEE

STE FEE

RTU

SWA-EAS FEE

RPW mags FEE

MAG FEE

HBIS Electronics Box

RTU

RPW ICU

RPW RAD

RPW PWS

RPW ants FEE

EUI MDE

EUI ICU

EUS MDE

EUS ICU

EUS Electronics Box

EUI Electronics Box

VIM MDE

VIM ICU

VIM Electronics Box SpW switch

DPU Core(LEON)

Payload Data Processing Unit

SpW router

SSMM

Routing Unit

SWA-HIS FEE

SWA-PAS FEE


NGD FEE

RTU

SWA-HIS FEE

SWA-PAS FEE


NGD FEE

RTU

STIX FEE

DUD FEE


SWA-EAS FEE

RTU

STIX FEE

DUD FEE


SWA-EAS FEE

RTU

COR MDE

COR FEE

COR Electronics Box

RTU

COR MDE

COR FEE

COR Electronics Box

RTU

RTU

RTU

RTU


P/F common

SIS FEE

HETn FEE

LET FEE

EPT FEE

STE FEE

RTU

SWA-EAS FEE


P/F common

SIS FEE

HETn FEE

LET FEE

EPT FEE

STE FEE

RTU

SWA-EAS FEE

RPW mags FEE

MAG FEE

HBIS Electronics Box

RTU

RPW ICU

RPW RAD

RPW PWS

RPW ants FEE

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

Standard CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

VIM

EUSStandard CV

EUIStandard CV

CORStandard CV

HBISStandard CV

HSISStandard CV

HSPISStandard CV


PDPUStandard CV

Standard CV

PDU

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

LC

L

PCDU

PDPU

Standard converter

LCL CV

Micro RTU

LCL

Space Wire

28 V regulated

On

/Off

com

man

dI/O,TM/TC

PCDU

P/L or P/L suite

PDPU

Standard converter

LCL CV

Micro RTU

LCL

Space Wire

28 V regulated

On

/Off

com

man

dI/O,TM/TC

Data management

Power distribution

Standard electrical interface


The scan platform for EPD sensors

Test

PDFE FPGA

IF

SRAMTEL 1

EPD-EPT

HVPS

TOF

I/F, Logic, Misc

SRAM

EPD-SIS

Start

StopPHA

Test

VLSIPHA

I/F, Logic, Misc

SRAM

EPD-LET

Test

VLSIPHA

I/F, Logic, Misc

SRAM

EPD-HET

EPD-STE

SSD/4

SSD/4

SSD/4

SSD/4

CSA/4

PHA/4Detector

I/FFPGA

Sensor common

LVPS

I/F to fixed part

EPD MISC CPUSRAM,

EEPROM

HK ADC

SSD BIAS

SWA-EAS

HVPSHVPS

HVPSHVPS

DetectorI/F

FPGA

Counter

SpW micro RTU

SpW with PDPU

Standard CV

28 V from S/C PCDULCLCV

P/F drive

electronic

P/F body fixed part

P/F mobile part

Test

PDFE FPGA

IF

SRAMTEL 1

EPD-EPT

HVPS

TOF

I/F, Logic, Misc

SRAM

EPD-SIS

Start

StopPHA

Test

VLSIPHA

I/F, Logic, Misc

SRAM

EPD-LET

Test

VLSIPHA

I/F, Logic, Misc

SRAM

EPD-HET

EPD-STE

SSD/4

SSD/4

SSD/4

SSD/4

CSA/4

PHA/4Detector

I/FFPGA

Sensor common

LVPS

I/F to fixed part

EPD MISC CPUSRAM,

EEPROM

HK ADC

SSD BIAS

SWA-EAS

HVPSHVPS

HVPSHVPS

DetectorI/F

FPGA

Counter

SpW micro RTU

SpW with PDPU

Standard CV

28 V from S/C PCDULCLCV

P/F drive

electronic

P/F body fixed part

P/F mobile part

Encoder

Bearings

Slip ring

Motor drive

P/F structure

Data, C&C Power

100 mm

100

mmEncoder

Bearings

Slip ring

Motor drive

P/F structure

Data, C&C Power

100 mm

100

mm

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 B


Overall payload mass budgetPAYLOAD 150,0

VIM 1 Optics 22,0 20% 26,41 Electronics 3,0 20% 3,6

EUS 1 Optics 6,7 20% 8,01 Electronics 6,0 20% 7,2

EUI 1 HRI (3) 4,6 20% 5,51 FSI 4,6 20% 5,51 Electronics 3,0 20% 3,6

COR 1 Optics 29,9 20% 35,91 Electronics 4,0 20% 4,8

STIX 1 Electronics 4,0 20% 4,8Delta to spec 0,0SWA-PAS 1 Sensor 3,0 10% 3,3SWA-HIS 1 Sensor 8,0 10% 8,8SWA-EAS 2 Sensor 1,5 10% 3,3RPW 1 Serach coil magnetometer 0,5 10% 0,6

1 Fluxgate magnetometer 0,5 10% 0,63 Electric antennas 1,6 10% 5,31 Electronics 5,5 10% 6,1

CRS 1 Electronics 0,2 5% 0,2MAG 2 Sensor 0,3 10% 0,6

1 Electronics 1,5 10% 1,7EPD-STE 1 Sensor 0,4 10% 0,4EPD-SPT 1 Sensor 1,3 10% 1,5EPD-SIS 1 Sensor 1,5 10% 1,6EPD-LET 1 Sensor 0,7 10% 0,7EPD-HETn 1 Sensor 2,0 10% 2,2EPD-DPU/H-L-VPS 1 Electronics 2,9 10% 2,6DUD 2 Electronics 0,5 20% 1,2NGD 1 Sensor 3,2 10% 3,5

1 Electronics 0,6 10% 0,7DPD 1 Electronics 2,7 10% 3,0NPD 1 Sensor 1,1 10% 1,2

1 Electronics 0,2 10% 0,2RAD 1 Electronics 6,1 20% 7,3Payload support elements 11,5

Boom 1 2,0 20% 2,4EUS filter 1 0,5 20% 0,6VIM filter 1 0,6 20% 0,7Scanning platform 1 Mechanism and structure 2,000 20% 2,4HOIS electronique 1 Electronique 3,000 20% 3,6HBIS elec 1 Electronics 1,500 20% 1,8


Instrument assessmentThe Unionics assessment


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.– Can be used as a pre-processor and dedicated interface to DSP21020 MCM.– Proven IPs are now available: “Space Wire”, 1553, ...etc)


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.


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.

– NOTE: A distributed system where data processing and compression is done at the payload level may reduce the interconnect data rate requirements, but it will not have any of the advantages of the centralised DPU unit approach listed above.


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, >200Mbit/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)”– Complete set of redundant interconnects (doubling the required number of

cables and connectors).


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) and Low mass (1.2kg).– Dual power rail design, requiring 2.5 and 3.3V +-10% regulated supply– Low power consumption:

– Standby (scrubbing and refreshing only): 1.5W per bank.– Simultaneous write and read at 16MHz: 3W per bank. – Simultaneous write and read at 40MHz: 5W per bank.

Disadvantages:– Works as a "tape recorder" storing and retrieving data serially, no random access.– The chip set for MM control require modification to include random data access capabilities.– Current design does not have a SpW interface.– Uses commercial 64Mx 8bit SDRAMs in non-hermetically sealed plastic packages.

(The above two-die per package SDRAMs have only been space qualified for LEO)


Assumptions for SOP payload DPU (Cont.) DSP21020 MCM building blocks

Advantages:– At least two different types of compact space qualified MCMs are available.– MCM designed with modular architecture in mind.– The Astrium MCM has built in three SpW I/Fs.– The 3Dplus MCM built in all the necessary program PROM.– UNIONICS software developed and based on DSP21020.– Space flight heritage on ESA space programs.– Low mass.

Disadvantages:– MIPs/W is not as high as other more recently available processors.– MIPs/g may not be as high as other more recently available processors.– Astrium MCM require external boot and program ROM.– 3Dplus MCM require external SpW I/Fs adapter (can be modified – NRE cost).


DS21020 MCM Options

Astrium DSP21020 MCM

Advantages:– Uses TSC21020F DSP, fully compatible with the Analogue Devices ADSP21020.– Two memory banks for program and data (128k x 32-bit SRAM each).– A processor peripheral controller ASIC. – A 1355 protocol controller ASIC, driving 3xSpW I/Fs through a 8k x 32 DPRAM.– Compact packaging design, 65g, 100 x 61 x 6mm.– Solder-less (Interposer) interface between MCM and host assembly.– A module design which can be populated with up to 8 x MCMs is available.– A module populated with 4xMCM and 8 way SWM is used within Inmarsat4 DSP.

Disadvantages:– Operating speed limited to 14MHz.– Require external boot ROM and program EEPROM.– Require external glue logic for boot and program load.– May require external shared RAM.


Astrium DSP 21020 MCM and Module

Generic 8 MCM Module (4 MCMs on each side)

DSP MCM


DSP21020 MCM Options

3D-Plus DSP21020 MCM Cube

Advantages:– Uses TSC21020F DSP, fully compatible with the Analogue Devices ADSP21020.– Two memory banks for program and data (128k x 32-bit SRAM each)– Three shared memory banks (512k x 32-bit SRAM each).– 4Mbit FLASH and 2K x 48-bit PROM.– A processor peripheral controller ASIC. – Compact packaging design, 200g, 52 x 52 x 33 mm.– Incorporated into ROSETTA, Mars’Epress and SMART-1.

Disadvantages:– Operating speed limited to 20MHz.– Require external external SpW I/Fs adapter (can be modified – NRE cost).– Pin Grid array connection to host board, may requiring mechanical support.– Require a heat-sink arrangements.


3D-Plus DSP 21020 MCM Cube

DSP MCM Cube

Typical implementation of a single cube module


Overview of Payload DPU Architecture

DSP Array Connection to 16 DSP MCMs.

– Z=16 x DSP21020 MCMs.– Each Each DSP has at least two SpW I/Fs.– DSP MCM is can be individually switched on/off.– Any DSP can be assigned as master controller.– Payload processing requirements can be met using Y<15 DSPs.– Z for Y redundancy arrangement, with N= Z-Y cold spared.

SpW I/F Nodes’ Array Connection to 36 prime and 36 redundant SpW I/Fs nodes.

– 34 prime and 34 (cold) redundant external Payload Instrument nodes.– 1 prime and 1 (cold) redundant external Spacecraft nodes.– 1 prime and 1 (cold) redundant mass memory nodes.– No single SpW I/F failure will reduce interconnect capacity.

Mass Memory Array 6 x 50Gbits Independent SDRAM blocks.

– Blocks connected together using proprietary internal reliable bus.– Common external SpW I/Fs, one prime and one redundant.– Each block can be individually switched on / off.– Only one block is required to be in active mode while others can be in standby.(This reduces the estimated total power as it assume a worst case of all blocks active)


Overview of Payload DPU Architecture (Cont.) SpW Switching Matrix Array

Implemented using 22 x 8 way SMX devices, split into four sections.– 8 x SMXs for prime SpW I/Fs nodes’ Array.– 8 x SMXs for redundant SpW I/F nodes’ Array.– 4 x SMXs for DSP array.– 2 x SMXs for interllink between SpW I/F nodes Array and DSP array.– Each SMX device can be individually switched on/off.

DSP Array SpW interconnect.– DSPs grouped into 4 groups each of which is associated with a single SMX.– 4 additional liks are provided directly between 8 of the DSPs to improve reliability.– Reliability of DSP SWM network is such that failure of any one SMX will not lose

more than 50% of the SpW link capacity to no more than two DSP nodes.

SpW I/F Nodes’ Array SpW interconnect.– SpW Nodes’ I/Fs grouped into 6 groups of 6 I/Fs each associcated with one SMX.– The 6 SWM connect to 2 SWMs which provide 4 SpW links.– An Identical SMX array is used to connect the redundant SpW I/F nodes.

Interlink SpW SMX.– Two (one prime and one redundant) SMX.– The prime and redundant SWMs connects to the prime and redundant SpW I/F

nodes’ SWMs respectively.– Each SWM connects to all four DSP SWM.


Data Processing Unit (DPU)

SMX

68 Ports SpW Switching Matrix

Payload1

Payload2

Payload34

DSP1

DSP2

DSP16

Spacecraft

Mass Memory (MM)APS

3 meters

Overview of Payload DPU Architecture


Key

8-Way SpW Switching Matrix

DSP21020 MCM with 2 SpW I/F’s

SpW I/F of the nth DSP MCM

n DSP MCM number

SpW Link

External SpW Connection

16

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

SMX

15

14

13

12

11

10

9

8

7

6

5

4

3

2

15 9

10

13

1

14

2 6

PRIME

REDUNDANT

SMX

n

n

SMX

SMX

SMX

SMX

SMX for Data Processing Unit


AVAILABLE SPECIFIC ACTIVE OFF16 10 8 2 DSPs 2 SPARE DSPs16 10 8 2 DPU SUPPORT LOGIC72 64 32 32 SpW TERMINAL LOGIC (assume 2 for 1 cold spared)23 20 13 7 SpW MATRIX includes > 1 S/C I/Fs (prime&redundant)34 30 30 0 PYLOADS 1 MM I/Fs (prime&redundant)6 5 4 1 MM UNITS 1 SPARE MM UNITS

MAX MAX MIN

MAX SPEC ACTIVE96.0 60.0 48.0 W ACTIVE DSPs8.0 5.0 4.0 W ACTIVE SpW5.8 5.0 3.3 W ACTIVE SpW MATRIX18.0 15.0 12.0 W ACTIVE MM127.8 85.0 67.3 W TOTAL DPU+MM POWER

36.0 32.0 16.0 W ACTIVE SpW TERMINAL LOGIC63.9 42.5 33.6 W APS DISSIPATION

191.6 127.5 100.9 W TOTAL DPU+MM+APS POWER126.1 W INCLUDING 25 % MARGIN

MAX SPEC MAX SPEC3.2 2.0 DSPs3.2 2.0 DPU SUPPORT LOGIC3.5 3.0 SpW MATRIX10.3 9.0 DPU+DSPs CASING + CONNECTORS

20.1 16.0 TOTAL DPU UNIT MASS18.8 16.7 SpW HARNESS AND CONNECTORS21.8 19.4 SpW TERMINAL LOGIC + CONNECTORS3.6 3.0 MM UNITS 0.5 mm DSP-MCM case9.6 6.4 APS 2 mm DPU case

53.8 45.5 TOTAL SUPPORT UNITS MASS73.9 61.5 kg TOTAL DPU+MM+APS+HARNESS+PAYLOAD I/Fs MASS

76.9 kg INCLUDING 25 % MARGIN

MASS

SPECIFICATIONS

POWER

SOP DPU Mass and Power Budget


DPU PROC. RATE 160 MIPS UNITS@ 20 MIPS at 20 MHzMM CAPACITY 200 Gbits UNITS@ 50 GBITS at 16 MHzx14bit/s=# PAYLOADS 30 UNITS 160 MINIMUM TOTAL REQUIRED MIPS INDEPENDENT OF PAYLOADS224 Mbit/s access rate

155.4 MAX TOTAL REQUIRED MIPS BASED ON NUMBER OF PAYLOADS

6 W DSP-MCM DSP MCM Standby Power 1.20 W dc0.5 W DPU SUPPORT LOGIC DSP MCM Power / MHz 0.24 W/MHz0.5 W SpW TERMINAL LOGIC0.25 W SpW MATRIX MM Unit Standby Power 1.67 W dc

3 W MM UNIT OF 50Gbits MM Power / MHz 0.08 W/MHz75 % APS EFFICIENCY

3 m SpW max cable length

27 g/cm3 Density of Al85 g/m SpW cable3 g/DSP-MCM 9 way MDM

200 g/DSP-MCM DSP-MCM200 g/DSP-MCM DPU-MCM SUPPORT LOGIC300 g/payload SpW TERMINAL LOGIC150 g/switch SpW MATRIX600 g/50Gbit MM50 g/W APS

POWER PARAMETERS

MASS

PROCESSING REQUIREMENTS

SOP DPU Mass and Power Budget (continued)


Study conclusions


Conclusion

Study demonstrates that 150 kg payload can be achieved– Pending size reduction of remote sensing instruments and key effect on mass & thermal– Thanks to relaxation of resolution to 150 km @ 0.21 AU– Size reduction and mass containment pave the way for the shortest cruise «ESA Flexible» mission

The main system issue is the management of the data volume– Manageable at spacecraft level– Critical at system level due to space to ground telemetry bottleneck

Study allows to clearly highlight and recommend– Instruments interfaces for accommodation on spacecraft– Instruments issues resulting from overall system environment

Study should be seen as a way – To get a common understanding ESA/science team/industry of

· Instruments requirements · Mission, environment constraints

– To better prepare spacecraft and system design· Avoid overdesign· Issue required level of interface information

integrated science payload for the solar orbiter mission final review

Documents

system studies

system assessment study

system level iterations

payload mass assumption

datestudy challenges

mass specificationallows

mass budget

instruments size