Spacecraft development - Lund Observatory · Spacecraft development Gaia program history Program...

Spacecraft development

Gaia program historyProgram overviewPayload optimisationProject organisationIndustrial team build upSubcontractor selectionOverall developmentSpecification treeAssembly, Integration and verificationAvionic modelMechanical and Electrical integrationTests

Gaia: a long story

Gaia is in the air since astrometry from space has demonstrated its viability with Hipparcos (1989 launch)

– Roemer mission by Hoeg & al in 1992

Industry is involved in Gaia definition from the beginning– First industrial studies funded by ESA around mid 90’s

– Nearly 10 years of advanced studies needed to issue a concept• Fulfilling as much as possible the science needs

• Technically and programmatically affordable

– Final concept settled in 2004

– Detailed definition and development started in 2006.

Gaia is to be ready for launch in 2011

Gaia programme overview

1995 – 2004: Phases O/A: Feasibility and Definition studies– Designing a valuable and affordable mission

– Iterations « as long as necessary » by a small industrial team, GST and ESA

– Technological support studies (TDA) launched in parallel of system studies

2005: Industrial competition to design & develop the spacecraft– Intensive period for

• ESA to issue Invitation to Tender

• Then for industry to reach a design/development maturity sufficient to commit on performances, cost and schedule

• And finally for ESA to select a Prime industrial

2006-2007: Phase B: Detailed design and team build up– Spacecraft design is tentatively frozen, building blocks are specified

– The industrial team is built up through competitive process

– The spacecraft performances are refined, with feed back form industrial team

– The Payload is optimised with GST support

2008 – 2011: Phase C/D: The spacecraft is developed, assembled, tested and qualified before delivery to ESA

Payload Optimisation with GST support

Although the schedule for the industrial team build-up and for the satellite development is extremely tight, Payload parameters are still optimised during phase B with the support of GST, after System Requirement Review

Main topics that deserved some optimisation– Photometric instrument: prism dispersions, red and blue band

definition,

– Radial velocity instrument: spectral resolution, star density, HR/LR modes,

– Windowing, sampling strategy, on-board processing algorithms

– Data priority management

– Scenario for scan law & modified scan law

Prime project organisation

spacecraft architects & coordinators

Mechanical & thermal engineering support Electrical & dynamics engineering support

Secretary & Documentation : C. Deblock

Configuration : JJ. Bouisset

Contract officer: J.A. VatinelSubcontracts: S. Peden

Project control: M. Mamy / I. BraultSchedule control: C. Buge

Product Assurance APM: D. Herbin

Industrial Devel’nt& AIT APM

M. Pendaries

System engineering& verification APM

F. Faye

Payload ModuleAPM

P. Charvet

CSWAPM

P. Humbert

Mechanical packageprocurement APM

C. Lebranchu

• Mechanical, thermal & propulsion C. Lebranchu• Launcher interface C. Lebranchu• Electrical systems & EMC G. Rougier• Performance verification & calibration E. Ecale• Dynamics & pointing E. Ecale

• Operations & Database X. Moisson• System datation X. Moisson• Data management & On-board software P. Humbert• AIT & GSE architecture engineering C. Gabilan

• Mechanical engineering F. Vogel• Thermal engineering U. Rauscher

• Electrical, RF & EMC enginnering A. Dyne• Dynamics & pointing enginneering T. Colegrove

Support to ESA for Science & Mission Performance expert: F. Safa

Electrical packageprocurement APM

P. Lelong

GAIA SpacecraftProject ManagerV. Poinsignon

spacecraft architects & coordinators

Mechanical & thermal engineering support Electrical & dynamics engineering support

Secretary & Documentation : C. Deblock

Configuration : JJ. Bouisset





M. Pendaries


F. Faye

Payload ModuleAPM

P. Charvet

CSWAPM

P. Humbert


C. Lebranchu







P. Lelong






M. Pendaries


F. Faye

Payload ModuleAPM

P. Charvet

CSWAPM

P. Humbert


C. Lebranchu







P. Lelong


Industrial team build-up

Right from the beginning of phase B2, the Core team is in place to organise the selection of the subcontractors for the complementary activities

Within the Core team, the responsibilities are assigned for the establishment of the rest of the industrial team

The list of items to be subcontracted is defined – more than 80 items

All fields of work are covered– Support tasks, AIT, sub-systems, equipments, software, GSE’s

The method of procurement– Open competition

The selection process– ESA Best Practices

– Confidentiality

– Fairness of competition

The industrial team build-up is a continuous process all along phase B2– Gradual release of ITTs

– Selection of subcontractorsmade by ESA and the spacecraft prime contractor on technical, cost, schedule, risk, and geo-return grounds

At the end of phase B2, the industrial team is complete and at work

Overall development

Development cycles

The specification tree

MRDAIV

requirementsPA

requirementsManagementrequirements

OIRDSGICDvol 1

SGICDvol 2

SRS GDIR

PLM spec E-SVM spec MSVM spec

ROOTSW URD

SYSTEMSW URD

PLM SWURD

AOCSSW URD

MMUSW URD

DMS SWURD

OperabilitySW URD

FPAspec

MPSspec

CPSspec

DSAspec

ESVMeqptspec

MSVMeqpt spec

MPS SWURD

TCS SWURD

FPAeqptspec

PLMeqptspec

CDMS, Power,TTC SW URD

GenericTMTC ICD

PL TCSSW URD

Doors links

Doors links

Doors links

Doors links

Doors links

Level 1 (ESA documents)

Level 2 (Spacecraft and SW URDs approv

Level 2 (Modules - same level as S/C but

Level 3 (Subsystems)

Level 4 (Equipments)

Assembly, Integration and Verification

The overall development covers– Design/ Development/ Manufacturing/ Assembly/ Integration/ Validation

– At unit, subsystem, module and spacecraft level

Almost all project people contribute to verification, to demonstrate that the spacecraft will fulfil its mission and meet the associated performance specifications – System and subsystem engineering teams

– Equipment unit procurement teams

– Product assurance teams

– Assembly, integration and test teams

– Instrument teams, for instrument, spacecraft, and system performance level

Verification can be performed by – Design Review

– Analyses

– Simulations or tests (or any combination of the above)• involving software models or physical development models

Verification by analysis or simulation

Ex: Pointing performances

Ex: Number of transit over the celestial sphere over 5 years

Avionics Model AVM

Mars ExpressAstrium Toulouse,

2001 to 2003

Gaia AVMAll electronic units engineering models on a table for functional tests

Mechanical and electrical integration

Mars ExpressAlenia Turin,

Jan-Sept 2002

Gaia assembly trees

Mechanical environment tests

Mars ExpressSine tests, Intespace, Dec 2002

Acoustic tests, Intespace, Jan 2003

Thermal environment tests, in vacuum chamber

PUMPING COOLING

RECOVERY

Thermal Validation

BalanceThermal

ValidationBalance Heating up Plateau Cooling down Plateau

Depressurise & set-up the initial

condition

TMM transient validation

OPS COLD

TMM Correlation OPS HOT

TMM Correlation Safe Mode

Heater validation.Coldest

Environment.

Set up the TV HOT

environment transient

Functional Tests at high temperature

Set up the TV COLD

environment transient

Functional Tests at low temperature

Return to Ambiant Press.

& Temp.

10130 hPa 10130 hPa CHAMBER PRESSURE < 10-5 hPa < 10-5 hPa < 10-5 hPa < 10-5 hPa < 10-5 hPa < 10-5 hPa < 10-5 hPa < 10-5 hPa

SOLAR FLUX OFF OFF OFF OFF OFF OFF OFF OFF OFF OFF

TTA & SA DUMMY Simulated Simulated Simulated Simulated Simulated Simulated Simulated SimulatedTEMPERATURE Not controled solar flux solar flux solar flux solar flux solar flux solar flux solar flux solar flux Not controled

(1280 W/m²) (1400 W/m²) (1280 W/m²) (1280 W/m²) (1430 W/m²) (1430 W/m²) (1100 W/m²) (1100 W/m²)

SVMTEMPERATURE Ambient (20°C) Ambient (20°C)

PROFILE

CONFIGURATION S/C LAM FOM (LGA) SOM (PAA) Specific Specific LAMAOCS SBM IGM (CPS) NM (MPS) thermal conf thermal conf SBM

PDHS OFF ON ON OFF OFF OFFCDU OFF ON OFF OFF

PLM Ambient (20°C) Ambient (20°C)TEMPERATURE

PROFILE-120°C

Specific ISST Specific ISST Specific ISST Specific ISST

OBJECTIVE

OPERATION

PHASE

TITRE

TB

SAFE COLDHOT

PLATEAUCOLD

PLATEAU

TV

Sub Phases

Mars ExpressThermal vacuum

with Sun simulation Intespace, Oct 2002

Gaia spacecraft test set up and

sequence

Gaia Payload Module in cryochamber in Liege

EMC environment tests

Conducted EMC, S/C openedTorino, Aug 2002

Marsis radar antenna Mock up characterisation,Intespace, Nov. 2002

Radiated EMC & RFCFlight configuration,Intespace, Jan 2003

Gaia,RF Auto compatibility

Gaia,/SoyuzRF Compatibility

Physical measurements

Mars ExpressMass, CoG, InertiasIntespace, Jan 2003

Gaia on its horizontal lifting device

Deployment tests

Beagle 2 QM ejection, Intespace, Nov 2002

Gaia DSA deployment ring

Gaia DSA mock up (ESA TDA)

Satelliteunder test

End to EndStimulator

PLMSVM

CDMU + EIU PDHU VPUTRSP/RFDU

PAA

CDUClks

End toend

FPAopticalstimul.

7 SpWSpW 7 SpWTC/TM HK+S

CDUEGSE

RF EGSE

RF

TM

S +

HK

TM S

PLM EGSEn°2

1

End to

end

OpticalstimulEGSE

Video TMHK + S

FPA

UmbilicalEGSE

TM/TC EGSE Trigger

Video TMHK + S

Video TM S

Dynamic FPAsimulator

- Sky simulation - SC motion simulation

Sky sim.files

7 S

pW

Scanning star simulator used for the GAIA FPA Technology Development

Test set up

From star signal acquisition until RF telemetry to ground

End to end functional test

Spacecraft design

Design approachMechanicalPropulsionThermalFunctionalDynamic and controlElectricalCommunicationsSoftwareFDIRTimingData Management

Spacecraft design approach

Although Gaia is unique by essence, the spacecraft design approach remains classical

No universal rules, but guidelines– Analysis and understanding of system requirements

– Split in elementary spacecraft functions

– Identification of critical items

– Reference to past experience

– Identification of system design option to assess (trade off approach)

– Favour the simplest solutions when applicable

– Consider « non technical » constraints in the background

• Cost / Schedule

• Geo return

– Iterative approach

Initial spacecraft sizing

Mission budgetary envelop and political constraints define the potential launcher

– ESA science missions: Soyuz Fregat is the current workhorse

Achievable spacecraft dry masses are deduced from launcher capability and required orbit manoeuvres– Soyuz-Fregat: 2.1 tons in direct insertion toward L2

– Insertion around L2 (Lissajous orbit) • 150 to 200 m/s needed for Gaia

First element to draw out is the instrument

Then preliminary allocation is made for other spacecraft functions

– A first idea of a possible spacecraft configuration is needed as soon as possible.

– It allows to get a first idea on system budgets and critical items

Spacecraft configuration build up

One of the very first activity in spacecraft conception is to draw out a configuration, with:

– An instrumental concept

– Overall shape, size and geometry (thus mass)

– Articulation concept

– Field of view, aerials and appendages • Payload instruments first

• Then antennas, thrusters, solar arrays, attitude sensors

Then a structural concept is issued– Interfacing with launcher (volume, physical connections and load path)

– Supporting the instrument

– Housing all servitude functions

And a 3-D model of the spacecraft pops up with all its elements– Necessary to evaluate mission performances and budgets

Mechanical architecture

Design approach1st step: identification of sizing elements• Support the large instrument

• Support the service electronics, the propulsion systems, the solar arrays

• Compatibility with launcher– Class of launcher defines the allowable volume and mass for the spacecraft

2nd step: Defining a mechanical architecture• Looking for instrument isostatic mounting to avoid distortion

• Looking for symmetry to avoid distortion

• Providing efficient load paths during launch

• Pre-sizing the structure (decoupling 1st eigenfrequencies wrt. launcher)

• Verifying balancing on the launcher

3rd step: specifying mechanical environment for units• Worst case worst place for everybody to initiate the process

• Conditions refined/decreased once accommodation exercise completed

Gaia mechanical architecture

Decoupling payload module/service module

Interface with launcher via a ring and a clamp band system

Central cone as main load path

Central Cone

Propulsion system

Design approach1st step: staging optimisation between launcher and spacecraft

» Gaia: launcher ensures the insertion into transfer orbit to L2

» Lisa Pathfinder: spacecraft/cruise stage composite inserted on a transfer orbit

» Bepi Colombo: multi stage spacecraft for Earth escape, cruise and insertion around Mercury

2nd step: selection of propulsion system• Analyse of orbit manœuvres and associated velocity increments

• Selection of the propulsion techniques» High thrust/medium efficiency (chemical) or low thrust/high efficiency (electrical)

» Cold gas, liquid or solid

» For the liquid system, selection between Hydrazine (Isp 220 s), bipropellant (Isp 300 s) or dual mode

3rd step: Propulsion architecture definition• Tank sizing• Thrusters configuration and capacity (in connection with AOCS)• Pressurisation and isolation circuit definition

4th step: Propulsion thermal control definition

Gaia chemical (NTO/MMH) propulsion system

Thermal architecture

Design approach1st step• Characterisation of the various mission phases

• Characterisation of the space environment through these different phases– Viewing factors wrt. Sun and planets

2nd step• Definition of sink temperatures

• Heat rejection budget for each spacecraft faces through the different phases

3rd step • Internal layout accounted for

• Thermal path definitions (conducted or radiated, no convection)

4th step• Architecture definition, control principle elaboration, material selection

• Radiators, heaters and heating power sizing

• Temperature map prediction

Gaia thermal architecture

DSA stowed and deployed

Functional architecture

Design approach1st step: centralised or decentralised • Weight of tradition (ESA science programs decentralised until Gaia)

• SW flexibility, CPU loads, modes and operations complexity

• Failure tolerance, hot/warm/cold redundancy

• Mass, power constraints

2nd step: looking for heritage• Mars Express: Rosetta, Mars Express, Venus Express, Telecom, Earth observation

3rd step: traffic and link budget establishment• Evaluate volume of data exchanged (on board, with ground)

• Define physical supports (data bus, point to point links)

• List the interfaces, standardise as far as possible

4th step: establish performance budgets• Define protocols• Establish CPU, memory and bus occupation budgets

Gaia functional architecture

CDU

Solar arrays

DeployableSunshieldAssembly

Chemical propulsion

STRE

Phased ArrayAntenna

SSPA

SSPA

SSPA

PowerConditioning

&Distribution

Unit

Battery

HeatersThermalsensors

VPU

VPU

VPU

VPU

VPU

VPU

VPU

VPU

VPU

VPU

PayloadData

HandlingUnit

28 Vtousers

VPU

VPU

VPU

VPU

VPU

VPU

VPU

Focal Plane Assembly

Payload Module

Electrical SVM

Mechanical SVM

Micropropulsion

ClockDistribution

Unit

SVM MIL-STD 1553B bus

Payload MIL-STD 1553 bus

SpW

SpW

SpW

SpW

SpW

SpW

SpW

SpW SpW

Tim

ing

Tim

ing

Pyropulses

STRE

I/O’s

STRESTRE

MicroPropuslionAssembly

BAM & WFSsource

electronics

M2 mirrorsmechanism

drive electronics

BasicAngle

Monitor

WaveFront

Sensor

M2mech

M’2mech

BAM & WFSsource

electronics

M2 mirrorsmechanism

drive electronics

ElectricalInterface

Unit

GyrosGyros

TRSP

X-Rx

X-Tx

TransponderTransponder

X-Rx

X-Tx

FineSun

Sensor

FineSun

Sensor

FineSun

Sensor

StarTracker

Gyros

STRE

MicroPropulsion

Elect.

LVDS

Structure SVM

SVM ThermalPassive Hardware

Thermal tent

Structure optical assembly

PLM Thermal Passive Hardware

Launch bipods

PkW

Control andData

ManagementUnit

Low GainAntennasystem

Electrical SVM units

CDMS units

AOCS units

Electrical power units

TT&C units

Redundant unit

PLM units

PLM units on PLM

LEGEND

Internally redunded unit

PLM units on SVM

Mechanical SVM units

Micropropulsion units

Chemical propulsion units

Solar arrays

DSA, thermal items

Links

Timing & datation

SVM MIL-STD 1553 bus

PLM MIL-STD 1553 bus

Space Wire/Packet Wire

RF signal

Dynamic and control

Design approach1st step• Considering the required attitude profiles and disturbing torque assessments:

– Selection of control concept: passive or active, spinned or 3 axis controlled

– Gaia: slow rotation with precession: 3 axes control

• Considering the spacecraft configuration and the disturbing torque assessment– Selection of the control (thrusters, micro thrusters, reaction wheels, control momentum gyros…)

– Gaia: ultimate stability required: micro thrusters

2nd step• Selection of sensors• Actuator sizing and control bandwidth definition• Gaia

– Hybridizing instrument (stability) and star tracker (inertial attitude) in normal mode

– Fine sun sensor and gyros for initial acquisition and safe mode

– Hybridizing high accuracy gyros/ start tracker for intermediate mode.

3rd step• AOCS modes definition, associated control laws• Definition of failure modes and associated recovery procedures• Pointing budget elaboration, resources budgets (propellant, processing…)

Gaia AOCS

Gyros

Astro

Sun sensors

Star trackers

Micro propulsionChemical propulsion

Gaia AOCS

Electrical architecture

Design approach1st step: preliminary resource sizing• Estimate power requirement per phase

• Battery sizing» Energy needs (launch phase, eclipses, manœuvres, peak power….)

» Select battery technology (function of cycling, required capacity)

• Solar array sizing» Considering power profiles, battery charging, attitude wrt. Sun

» Select solar cell technology (deep space probe or not, temperature range)

2nd step: selection of a power topology• Conditioning and regulation principle

• Bus voltages

• Line protection

3rd step• Define electrical interfaces• Establish power budgets

Gaia power system

Li-ion battery

Solar array

Power Conditioning and Distribution Unit

• Triple Junction GaAs cells on solar arrays• Lithium ion batteries based on standard Sony cells• 28 V regulated and maximum power point tracker to minimise instability

CommunicationsDesign approach

1st step: Selection of frequency • Imposed by regulation agency (ITU),

• Function of mission type and orbit

• Gaia from L2: X band (8 GHz) TM/TC , 10 MHz max TM bandwidth

2nd step; Definition of communication scenarios• Gaia:

» Cebreros 35 m ESA station 8 hour/day outside galactic plane scans

» Cebreros and New Norcia 35 m ESA stations during galactic plane scans

3rd step: resources sizing (iterative)• Ground/spacecraft link capabilities: preliminary link budgets as function of distance, antenna

geometry, memory capability, data availability requirement, protocol

• Antenna, amplifiers, transponders selection in iteration with spacecraft configuration, power resources and thermal control capabilities

4th step: on board architecture definition: • Interconnection schemes

• Redundancy, reconfiguration,

• Detailed link budgets, system budgets

Gaia Communication system

Choked horn LGA

PAA radiating cone

Phased Array Antenna electronics

X band transponder

- X LGA-1

RF Splitter 2:7

PAA

TC

TM

TRSP-1

RTDiscreteTM/TC

X-BandRx

X-BandTx

RTDiscreteTM/TC

TC

TM

TRSP-2

X-BandRx

X-BandTx

TT&C

RT-A

DiscreteTM/TC

RT-B

DC/DCCV B

DC/DCCV A

EPIC A

10 Ms/s10Ms/s

Discrete

- X LGA-2

DC/DCCV

DC/DCCV

TM/TC

28V

+ X LGA

RFswitch

RFswitch

DiscreteTM/TC

SHPHK TM

BM 1

SSPA Qua

dri-M

odu

le 1

BM 2

SSPA Quad

ri-M

odu

le 2

BM 3

SSPA Qua

dri-M

odule 3

BM 4

SSPA Qua

dri-M

odu

le 4

BM 5

SSPA Q

uadri-M

odule 5

BM 6

SSPA Quad

ri-M

odu

le 6

BM 7

SSPA Quad

ri-M

odu

le 7

EPIC B

Controlfunction

B

Controlfunction

A

4ks/s

C

C

C

C

CC

CC

4ks/s

28V28V 28V

DiscreteTM/TC

DiscreteTM/TC

DiscreteTM/TC

SVM MIL-STD-1553B Bus N

SVM MIL-STD-1553B Bus R

Central software

In charge of spacecraft managementHW / SW interface, data acquisition and processing

AOCSSensors, Actuators, Control laws, mode management

Data managementData handling (services, mailboxes, priorities, events, interruptions,…)

Bus management (PLM and SVM 1553 buses, SpaceWire bus)

Mission Time Line

TM generation, TC acknowledge and processing, tasks scheduling

Dump / Patch

ThermalThermal control loops management

ElectricalLCL management, Pyro sequences

CommunicationsEarth ephemeris for PAA beam orientation

Monitoring Detect the failure

Isolate the suspected units

Reconfigure on safe configurations

Gaia software layered architecture

Failure Detection Isolation & Recovery

Hierarchical FDIR: the spacecraft detects autonomously on board failures and tries to continue its mission or place itself in safe mode through a gradual recovery process.

Groundcommands

Unf

ore

see

n e

ven

t

Level1

OK

OK

OK

OK

NO

K

NO

K

NO

K

NO

K

Redundantprocessor & units

Nominal Processor& redundant units

Redundant PM& nominal units

Last chanceconfiguration

Science continued

SpacecraftSafe Mode

(SSM):mission

interrupted

On boardMission Time Line

Spacecrafttelemetry

Gro

und

sta

tion

pa

ss

Nominal autonomy Nominal autonomyAutonomous hierarchical FDIR

Gro

und

sta

tion

pa

ss

Gro

und

sta

tion

pa

ss

Localreconfiguration

Level2

Level3

Level4

Specific on Gaia: Timing architecture

PPSRegister

OBTRegister

OBT

XO1553 I/F

SynchroUnit

PpsTo

CDMU

SCET

PEM 1 PEM #7

Time CodeGenerator

VPU #i

IM #i

TDI1

TDI2

TDI3

20 MHz

OBS

Latch signal

Time Strobe

TFG

MasterClock

Delivery of TDI and master clocks to the video chain, all of them being derived from the Hydrogen Maser (RACM)

Time correlation with ground time scale (UTC) thanks to Time strobe generated by the CDMU

PAA

Science data management architecture

A long way from Focal Plane to SOC and DPAC…

Kourou, French Guyana

Baikonur, Kazakhstan

Launch Campaign

Departure to launch site

The S/C shall fit in its container that shall fit within Antonov airplane

Final preparation

Propellant fueling

Installation on launcher, encapsulation, transport

Soyouz in Kourou(From Arianespace documents)

Babysitting on launch pad

Launch

And spacecraft operations are starting in ESOC

Gaia configuration: 1998

Science Mission Requirements

Two Astro viewing directions

Hipparcos scan law, sun-spin axis angle ξ > 45 deg,

Scan rate > 60 arcsec/s, precession period fixed @ 70 days

Astro detector operating temperature < 170 K

Astrometric accuracy ~ 10 µas for V > 15, limiting magnitude V ~ 20,

RVS spectral range 849 nm – 874 nm, resolution > 0.075 nm with a goal at 0.0375 nm per pixel

5 spectral broadband filters in astro fields (Broadband Photometer), 11 Medium Band filters in RVS telescope field (Medium Band Photometer)

Lissajous orbit about L2No eclipse for 5-6 years

Sun-Satellite-Earth angle < 10°

SUNSatelliteSpin axis

55°

Astrometric InstrumentLine of Sight 1

Astrometric InstrumentLine of Sight 2

Precession of the spinaxis in 72 days

GAIA

ConsecutiveGreat Circles

Basic Angle

Gaia configuration : 1998

Sunshield φ 9.5MSolar array(Rear face MLI covered) Spectrometric instrument

entrance aperture

Astrometric instrumentsentrance aperture

TT&C antenna

Star sensor

PLM Thermal cover

Optical cover

& baffle

Sunshield φ 9.5MSolar array(Rear face MLI covered) Spectrometric instrument

entrance aperture

Astrometric instrumentsentrance aperture

TT&C antenna

Star sensor

PLM Thermal cover

Optical cover

& baffle

Science telemetry

Antenna

TT&C antenna

10 N thruster

Solar array

Sunshield φ 9.5M

Propulsion subsystem

Science telemetry

Antenna

TT&C antenna

10 N thruster

Solar arraySolar array

Sunshield φ 9.5M

Propulsion subsystem

Satellite dry mass (incl. margin) = 2035 kgPropellant mass = 1010 kgTotal launch mass with margins = 3150 kgSatellite power budget (observation) = 2475 WLauncher: Ariane V

General Approach

Required performance

Optical quality: overall WFE < λ/14 rms

Basic angle variations/monitoring over 1 revolution (3 hours): ~ 1 µas

Minimum active control & mechanisms by taking full benefit of L2 orbit stable environment:

Sun shield/solar panels protects the payload from sun radiation.

Radiative & conductive de-coupling from SVM

Athermal design, by using of a single material, homogeneous & isotropic, for optics and major structural parts,

No mechanism in nm or pm range

Monolithic mirrors and refocusing mechanism (few µm accuracy).

Basic angle monitoring in orbit

Zerodur: Good optical properties, very low CTE @ RT (degrades in cold), but poor structural properties & lightweighting capability,

CFRP: Good structural properties & lightweighting capability but poor optical properties,

C-C (Aérospatiale data): Good structural properties, very low CTE but poor optical properties & low thermal conductivity

SiC: Good structural properties, very good lightweighting capability, low CTE (improves in cold, 0.5 @ 100 K) & good optical properties. Selected for Gaia design.

Parameter Zerodur Carbon-Carbon (C-C) Silicon Carbide (SiC) CFRPYoung Modulus E (GPa) 90 60 420 117Density ρ (g/cm3) 2.53 1.56 3.15 1.65Ultimate tens. strength (MPa) 20 190 350 240CTE α (µm/m/K) 0.05 -0.1 2 0.46Thermal conductivity λ (W/m/K) 1.5 7 190 37Lightweighting ratio E/ρ 35 38 130 71Thermal distortion ratio α/λ (x 100) 3.4 1.4 1 1Polishing yes no yes noCoating yes yes yes yes

Note: Values given at room temperature

Material selection

SiC Technology Maturity

Demonstrator telescopeΦ200 mm, WFE < 20 nm,Tested –100°C to +60°C

OSIRIS TMA telescope, 12 kgΦ 90 mm, WFE < 20 nm,

Trange -100°C to 70°C, will fly on Rosetta

Sofia secondary reflector, 1.9 kgΦ 352 mm, WFE < 50 nm

Rocsat2 telescope, 75 kgΦ 630 mm, WFE < 40 nm

Herschel demonstration reflectorΦ 1350 mm, made of 9 brazed segments

WFE < 4 µm, tested @ 110 K

Herschel Full scale petal, 1500 mm

SiC Technology Maturity

Herschel SiC telescope diameter is 3.5 m, mass 300 kg

Reflector and telescope structure are made of SiC

Operating temperature ~ 80 K

Delivery to ESA early 2005

All facilities used for Herschel are applicable to Gaia

Major features of the study design

Use of large aperture independent telescopes. No need for high accuracy mechanisms (nanometric positioning accuracy required for the case of an interferometer)

Common telescope for ARVI & narrow-band photometry (Arvi Photometer Telescope).

Feasible detectors, pixel dimensions > 9 µm,

Short term basic angle stability ensured by the use of Silicon Carbide for both the reflectors and the structure,

Implementation of a device capable of monitoring basic angle variations with an accuracy compatible with Gaia needs

ARVI (Auxiliary Radial Velocity Instrument) located on the symmetry axis

Design compatible with Ariane V launcher and ESA assigned cornerstone budget

Optical assembly configuration

ASTRO-1 secondary mirror

ASTRO-2 primary mirror

Basic anglemonitoring device

ASTRO-1 focal plane

ASTRO-1 primary mirror ASTRO-1 tertiary mirror

Optical assembly configuration

Astro Beam (1/4)

Astro Beam (2/4)

Astro Beam (3/4)

Astro Beam (4/4)

SPECTRO Tertiary Mirror

SPECTRO Secondary Mirror

SPECTRO Focal Plane

SPECTRO Primary Mirror

APT instrument

APT Beam (1/4)

APT Beam (2/4)

APT Beam (3/4)

APT Beam (4/4)

Dimensions compatible with Ariane 5 fairing

Payload accommodation

Optical bench design

The design of the torus is established in order to ease manufacturing and AIT activities

– the internal reinforcement will be chosen with respect to the equipment location in order to decrease the mass

150 mm300 mm

thickness 5 mm

GAIA Concept& Technology Study

Focal Plane & Electronics Accommodation

The Focal Plane Units are mounted on the PLM torus with a specific thermal insulation

The Video Processing Units are mounted on the SVM structure

FP active face

MLI +

permaglass wachers

MLI

torus

radiator

SPACEPLM CAVITY

SVM closure panel

torus

VPU

support +

radiator

Secondary structure design

The shape of the tent is optimised in order to stay under the protection of the sun shield and it has no connection with the PLM to improve stability

MLI

carbon truss

The tent is made of carbon truss supporting MLI on both faces

GAIA mechanical model

GAIA mechanical analysis

First lateral mode First axial mode

Electrical Architecture

FPAASTRO-1

FPASPECTRO

FPAASTRO-2

Video Processing Units (VPU)ASTRO-1

Video Processing Unit (VPU)RVS / NBP

Video Processing Units (VPU)ASTRO-2

Payload Data

Handling Elec.

(PDHE)

Solid State Recorder

(SSR)

High Rate Telemetry Formatter (HRTF)

Science TM Comm's

Link

PowerDistrib. Unit

Thermal Control & Mechan. Drive Elec

Central Data Management Unit

(CDMU)

Syncro /Time tag

PLM SVM

Science Data Acquisition / Processing / Transmission

SVM General Services

Payload equipment

Data

Data

Data

Comd / Contl Bus(1553, OBDH, ...)

1

© Astrium76

Gaia Configuration : 2002

PLM

SVM

Sunshield

PLM

SVM

Sunshield

Gaia configuration : 2002

PLM

SVM

Sunshield

PLM

SVM

Sunshield

Satellite dry mass (incl. margin) = 1150 kgPropellant mass = 250 kgTotal launch mass with margins = 1500 kgSatellite power budget (observation) = 1500 WLauncher: Soyuz Fregat

2002 Study concept and objectives

Study requirements & objectives, mainly to reduce the overall cost :– Soyuz-Fregat evaluation against Ariane 5 Reduce mass and size– Re-open basic trade-offs (Solar array / sunshield; payload telemetry antenna)– Update and Optimize of the development & validation approach– Update the technology planning– Optimize the industrial management approach

Study assumptions :– Same science requirements as for the Concept & Technology Study of 1998;– Re-use of up-to-date service modules strongly recommended– Transfer to L2 performed by launcher

GENERAL PARAMETERS

Previous Design New Design

Observing time L L = 4 years L = 5 years

Scan rate ω 120 arcsec/s 60 arcsec/s

Precession period ωp 70 days 70 days

Rotation axis 55° from sun direction 50° +/- 0.1° from sun

direction

Star population V < 20

Average value

« Worst case »

Ns = 14300 stars/deg²

155 000 stars/deg²

Same assumptions

Total number of observed stars ~ 1 billion ~ 1 billion

2002 Payload Optimisation

Reduce mass and size for being compatible with Soyouz launcher

Total launch mass ~ 2270 kg (A5) : mass objective < 1500 kg ( - 34%)

Reduce the PLM mass ~ 900 kg:Astro focal plane assemblies ~ 168 kg (with video electronics)

mass decrease obtained by reducing the number of CCDs

Optics with their mounts ~ 270 kg :

mass decrease is obtained by reducing the reflector sizes

Structure ~ 290 kg

structure mass decrease is a consequence of size and supported mass reduction

Mass reduction objective is in conflict with accuracy requirement.

Design improvement is mandatory for Soyouz compatibility.

Reduce overall complexity and cost– Try to harmonise CCD definitions

– Minimise hardware tests

– Suppression of mechanisms whenever possible

2002 Design Features

Astrometric Instrument– Superimpose the two focal planes, by making use of a focal plane combiner

located close to the exit pupil

– Increase the overall optical path length for practically suppressing optical distortion

– Enlarge CCDs along scan (~ 60 mm x 45 mm)

– Passive cooling of the FPA detectors down to 160 K – 170 K (radiation)

Spectrometric Instrument– Along scan dispersion for RVS

– Suppression of RVS mechanism

– RVS spectral resolution improved by a factor 2

– Medium Band Photometry improved

2002 Astrometric Instrument Configuration

M1

M2

M3

M4(1/2 focal plane

combiner)

M5, M6(common reflectors)

Astro focal plane

FPA radiator

6 mirrors in totalM4, M5 and M6 connected or common to both pathsFocal plane common to both paths

2002 Spectrometer baseline

Along scan disp. - Horizontal telescope Scale: 0.10 AsF 11-Mar-02

255.10 MM

Scan axis

Needs more optimisation to :Try to reduce diameters / mass / number of lenses

Get more realistic air gaps and lens thicknesses

M1

M2

RVS – MBP not shown

M3

13 lenses for the RTVS optic

2002 RVS Telescope accommodation

M1M2

MBP & RVSM3

2002 PLM mechanical analyses

A Finite Element Model of the PLM is constructed for a preliminary verification of the mechanical design

2002 FEM results: PLM eigenfrequencies

1st mode: f > 31 Hz (lateral) 1st axial mode (local): f = 54.5 Hz (lateral)

Frequency requirements on PLMf > 20 Hz, in lateral

f > 45 Hz, in longitudinal

Frequency requirements are met with a good margin

2002 FEM results: stress analysis

Stress levels are compatible with silicon carbide strength

Stress distribution 20g axial

Stress distribution 15g lateral

Stress analysis has been made for the following quasi static loads:

20 g in longitudinal direction

15 g in lateral direction.

PLM / SVM mechanical interface

The upper floor of SVM is a stiff and stable plate with following functions:

Support the in-plane forces due to the iso-static device of the optical bench

Transmit the lateral loads to the SVM central tube

Support the PLM thermal tent

Offer a high stability during the 6h period• No dissipating equipment on this plate• Thermal insulation from SVM and PLM• Stable CFRP structure

Finite Element Model

Rough model : used for a first estimate of stiffness and strength

Detailed model : used to analyse the global satellite behaviour

Modal analysis results

Lateral mode Y

Axial mode X

Lateral mode Z

Modal analysis results

Lateral mode3 Y=27.26 Hz

Lateral mode4 Z=27.56 Hz

Recommended technology developments

The sunshield TRP shall be oriented towards a dedicated system – without solar array – and assess the feasibility of deployment and its compatibility with the shield cover itself

The FPA electronics and video processing units must be studied on technical and industrial points of view. A dedicated TRP is recommended on these subjects.

The CCDs

The Micro propulsion technology selection

3

© Astrium92


Satellite dry mass (incl. margin) = 1400 kgPropellant mass (incl. Cold gas) = 280 kgTotal launch mass with margins = 1830 kgSatellite power budget (observation) = 1700 WLauncher: Soyuz Fregat


Key figures and order of magnitude

Primary mirrors: 1.45 m x 0.5 m

Focal length: 35 m

Pixel: 59 x 177 mas (10 μas 1/6000 pixel)

1 μas rotation M1 < 10 picometers at the edge

Focal plane: 420 x 850 mm

About 100 CCD, ~ 1 Gpixel

Star on CCD: mean: 150, peak: 36000 (magnitude 20)

Stellar flux: 20 000 e-/s @ V=15 200 e-/s @ V=20

Sample datation accuracy: 15 ns

Tore (3 m dia) thermal stability required some tens of μK

Rate measurement error < 0,9 mas/s

Rate pointing error < 5 mas

Attitude High Frequency Disturbance < 3.4 μas

2005 PLM design

M1

M3

M2

Common FPA

Common folding mirrors

M1

M3

M2

M1

M3

M2

M4 M’4

2 Three Mirror Anastigmat telescopes M1-M3

Intermediate image used for field discrimination

Beam combiner at their exit pupil => common FPA

Full silicon carbide architecture

Passive thermal design

Decoupling by release of launch bipods

2005 PLM design

RVSdetectors

Red & blue photometerdetectors

Astrometricfield

Sky mapper

BAM & WFS

M4/M’4beam combiner

Photometerprisms

RVS gratingand afocal

field corrector

M5 & M6fold mirrors

RVSdetectors

Red & blue photometerdetectors

Astrometricfield

Sky mapper

BAM & WFS

M4/M’4beam combiner

Photometerprisms

RVS gratingand afocal

field corrector

M5 & M6fold mirrors

Focal Plane Assembly

A cold part at 160K for limiting CCD sensitivity to radiations

A warm part for electronics

Bipods for mechanical mounting & thermal decoupling wrt optical bench

Photometer prisms

CCD radiator& shielding

Electronics radiatorFPA bipods

ElectronicsRadial VelocitySpectrometer CCDs

Astro FieldCCDs

Blue & RedPhotometer

CCDs

Star MapperCCDs

WFSOptics& CCD

BAM CCD

0.4

20

m

0.930 m

row 1

row 2

row 3

row 4

row 5

row 6

row 7

WF

S2

AF

1

AF

2

AF

3

AF

4

AF

5

AF

6

AF

7

AF

8

AF

9

BP

AS

M1

AS

M2

RP

WF

S1

BA

M-N

BA

M-R

RV

S1

RV

S2

RV

S3

Service Module Design

3

© Astrium98

Gaia Configuration: 2007

Gaia configuration evolution: 2007

Satellite dry mass (incl. margin) = 1620 kgPropellant mass (incl cold gas) = 350 kgTotal launch mass with margins = 2050 kgSatellite power budget (observatiuon) = 1750 WLauncher: Soyuz Fregat

Gaia: key figures and order of magnitude

Primary mirrors: 1.45 m x 0.5 m

Focal length: 35 m

Pixel: 59 x 177 mas (10 μas 1/6000 pixel)

1 μas rotation M1 < 10 picometers at the edge

Focal plane: 420 x 850 mm

About 100 CCD, ~ 1 Gpixel

Star on CCD: mean: 150, peak: 36000 (magnitude 20)

Stellar flux: 20 000 e-/s @ V=15 200 e-/s @ V=20

Sample datation accuracy: 50 ns

Tore (3 m diameter) thermal stability required some tens of μK

Rate measurement error < 0,9 mas/s

Rate pointing error < 5 mas

Attitude High Frequency Disturbance < 3.4 μas

S/C launch mass 2.1 tons

Solar Array capability 1.9 kW

Mass memory capability 1 Tb

S/C Height 3 m

Deployed Sunshield ø = 10 m

The solar array has fixed panels and deployed panels mounted on the sunshield

The CFRP SVM structure is mainly made of lower & upper floors, a central tube including the LV/IF ring, radial panels and a DSA ring

Bi-propellant propulsion system with 8 x 10 N thrusters for orbit correction

Cold gas micro-propulsion to perform stable attitude control

[1μN – 500μN] thrust range

The mechanical SVM

The harness provides electrical connection between units for power and data management

Thermal Tent Structure (TTS)

Thermal Tent Hardware (TTW)

Deployable Sunshield Assembly (DSA)

Solar Arrays

Structure

Chemical Propulsion System (CPS)

Micro Propulsion System (MPS)

Harness

The thermal tent provides a stable thermal environment and protects against micro-meteoroids ; it has large cut-outs for the instruments FoV and the FPA

The 10 m dia. sunshield is deployed in orbit to keep the PLM in the shade of the sun ; it has to be very flat and uniform, with an overall 0.2 deg flatness requirement

The electrical SVM

The electronic units are accommodated orthoradially

in the SVM cavitiesVery stable power dissipation and avoidance of conductive thermal paths minimises thermoelastic

distortion

AOCS3-axis controlStar trackers/gyro/sun sensorsASTRO instrument in the loop for accurate pointingThrusters and micro-thrusters for attitude control

X-band TT&CThree omni directional LGAPhased Array Antenna with up to 28 SSPA

– No mobile parts, electronically steered beam– TM only, 4.8 to 8.5 Mbps

Two X band TRSP

Power73 Ah Lithium batteryPower conditioning and distribution unit

– 28 V regulated main bus– MPPT regulation of 1.9 kW solar

arrays– Protected power lines to users

Data managementCommand & data management unit

– ERC 32 processor (N+R)– 8 Gb memory for mission management

Electrical Interface Unit– I/O customised with SVM and PLM units

The PLM

All SiC structure and mirrors for the optical bench, very high thermo-mechanical stability required for TDI and basic angle

A large focal plane with ~ 1 billion pixels.CCD passively cooled to 165 K

PLM electronic units accommodated in the SVM (PDHU, VPU, CDU)

Payload configuration

Torus design

FPA design

Satellite configuration evolution summary

A priori very different, but in fact implementing comparable principle

Thank you

Xavier [email protected]

Spacecraft development - Lund Observatory · Spacecraft development Gaia program history Program...

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