electrical power systems - uliege · 12 ref: vlr/080066 date : 28.01.08 introduction / tas-etca...

Ref : VLR/080066Date : 28.01.08

Electrical Power Systems

Vincent Lempereur

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Agenda

1. Introductionn Presentation of Thales Alenia Space ETCAn Presentation of myselfn EPS – general information

2. Primary power sourcesn Solar cells & solar arraysn Fuel cellsn RTGn Others

3. Secondary power sources - batteries4. Power Management, Control & Distributionn Architecturen PCU / PCDU

5. Power budget – practical exercise6. Conclusions

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Agenda

1.1. IntroductionIntroductionnn Presentation of ThalesPresentation of Thales AleniaAlenia Space ETCASpace ETCAnn Presentation of myselfPresentation of myselfnn EPS EPS –– general informationgeneral information




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Introduction / TAS-ETCA

World Leader in critical information services

Aeronautic Space

Aerospace 25%

Air systems

Terrestrial & inter-arm systems

Navy

Defense 50%

Solutions Security Services

Security 25%

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67 %

33 %

33 %

67 %

* 12 milliards eurosof turnover (2006)* 68 000 employees

* 12,5 milliards eurosof turnover (2006)* 58 000 employees

European Leader in solutions by satellitesEuropean Leader in solutions by satellitesLeading position for orbital infrastructuresLeading position for orbital infrastructures

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… Worldwide repartition

San FranciscoSan Francisco

ColombesColombes

AnversAnvers

ToulouseToulouseCannesCannes

RomeRomeL’AquilaL’AquilaFlorenceFlorence

MilanMilanTurinTurin

MadridMadrid

CharleroiCharleroi

USA Europe : 11 industrial plans

KourouKourou

French Guyana

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First European satellites manufacturer

In the fields of …

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More than 40 years of

Experience in Space

Thales Alenia

Space ETCA

1963 - 2006

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Thales Alenia Space ETCA 2007 key figures2007 key figures

CA : 75 millions €

Telecoms & Security

Space

Export: 75% of profits

+/- 568 people

23%77%

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Thales Alenia Space ETCA 2007 key figures2007 key figures

Workers 23%

Ingineers & management

44%

Employees33%

Sales & administration

11%

Engineering40%

Manufacturing & Quality

49%

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Business segments

§§LaunchersLaunchers §§Telecoms & SecurityTelecoms & Security§§SpaceSpace

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Space activity / Equipment and Services for Space Applicationsn Satellites

n Electrical Power treatment and distribution system• Geo telecom S/C

• Spacebus (2000, 3000 et 4000) from 4 to 20 kW• CAST (China)

• Earth observation and scientific missions• Proteus from 250 to 700 W• Mean Power from 750 to 2.5 kW• Microsat < 250 W

n Avionicsn Power supply for Plasmic Propulsion thrustern Power supplies for Traveling Wave Tubesn DC/DC Converters for electronic equipmentn Specific Check Out Equipment (SCOE)

n Launchersn On board Control and security Equipmentn Overall and Specific Check Out Equipment (OCOE/SCOE) n Electrical Ground Support Equipment

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Some Space ReferencesAriane 5, Soyouz, AMC12, AMC13, Apstar, Syracuse, Star One, HotBird 7, Herschel-Planck, Chinasat, Galileo, Express AM11-22, Koreasat, Smart 1, ATV, Expert, Ciel2, Turksat, W2A, W7, Satcom BW, Pléiades-HR, Spirale, Alsat-Elisa, Siral 2,Taranis-SMESE, Stereo, Globalstar …

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Launchers / on-board equipments

2 BCSand

2 BDP

2 BCSand

2 EPH

2 EPHand

2 BDP

2 CDC

2 BCSand

2 EPH

2 BCS

2 EPEor

1 EPSM

Largest supplier of electronic

equipment forAriane 5

•83 kg of electronics/LV

•21 devices/LV

•20% of the AR5 electronics

•75% of the equipment bayelectronics

•4 to 6 LV/year

CASE CASE

EPC

EPC

EAP1

EAP2

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Launchers / on-board equipments

TAS ETCA is prime for the safeguard system of Soyuz:

BCA

JA1

JM1

JT1

JF1

JC1

JS1

JF2

JB2

JB1RTC1J1

J4J3

RR1 J2J1

RR2 J2J1

RTC2J3

J4J1 JA2

JM2

Bat2

Bat1

JT2

JC2

JS2

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Launchers / check-out equipmentsn Ariane 1 to 4

n Check out of launcher electrical performancesn Application software development

n Ariane 5n Check out of Ariane 5 thrusters (4 equipment)n Ariane 5 launcher family (14 equipment)

n Equipment for satellites and sub-systemsn Meteosat Second Generation PDUn Plasmic Propulsion Power Unitn SPOT 5 RSJDn SB4000 PCU

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Manufacturing & testsn Building: 42 000 m²n Manufacturing & tests : 22 000 m²n Clean rooms : 10 000 m²

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Introduction / V. Lempereur

Trainingn Electrical engineer from ULG (1996)n MBA in IEFSI (Lille)

Professional career in TAS-ETCAn Designer on PCU & PPU projects

n SB4000 SBR & PCU, SB4000 PPU & FUn BCC on SOYOUZ for Globalstar1 project

n ETCA’s resident in Cannes during two years (EPS architecture / PCDU phase A / …)n Wales / Earthcare / Spectra / Exomars / Galileo / Pleiades / …

n Technical manager on PLEIADES DRU (PCDU)n Project manager on PCDU / PSU projects

n SIRAL PSU, ALTIKA, MIRIn µSAT PCDU

n PCDU/PCU product manager

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Introduction / EPS general information

A satellite is made of…

P/F (Platform)n Mechanical & thermal structuren Electrical system, avionic, propulsionn On-board computer, software, remote controln Energy sources: solar, batteries, fuel

P/L (Payload)n Antennas, TWTA, …n Camera, altimeter, radar, detectors, …n Clock, scientific instruments,…

P/FP/L

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Satellite Electrical Power Subsystem (EPS) shall

n provide to each S/C platform and payload equipment the required power over the whole mission

n have energy capacity to power equipments in case of orbital night phases, transient phases and peak power demand

n autonomously manage the available power in order to provide the equipment’s power and to charge the battery

n fulfill some distribution requirements providing ON/OFF protected power lines, heater supply (for S/C thermal control needs) and commanding pyro lines (e.g. SA and antenna deployment)

Note: power system failure means the loss of mission

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General functional block diagram

Primary Primary sourcessourcesPower

generation

SecondarySecondarysourcessourcesEnergy storage

Power managementPower managementRegulation & Continionning

Power Power distributiondistributionDistribution &

protection

User’sUser’sSpacecarfts

loads

Command & ControlCommand & ControlInterfacesInterfaces

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Functions (1)

n Power generationn The power is generated from different sources (‘fuel’) or combination of them: the

Solar radiant energy (solar cells via photovoltaic effect), Chemical (piles – fuel cells), nuclear (RTG), mechanical (reaction wheels), …

n Primary sources convert ‘fuel’ into electrical power

n Energy storagen The energy is generally stored under a electro-mechanical form and retrieved

under an electrical formn The storage of the energy is done by a secondary source, when the primary

system’s energy is not available or insufficient

n Conditioning and regulationn This function covers everything which is required to adapt the primary sources to

the need of users ‘equipment’n Regulators

• To maintain a constant voltage or current• Regulation of battery charge and discharge, regulation of the commutation of solar generator sections

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Functions (2)

n Distributionn To distribute the conditioned power to usersn DC/DC voltage convertersn ON/OFF switchesn Does not include the harness

n Protectionn To avoid a propagation of failures or any Single Point Failuren Protections against short-circuits

• Fuses• Circuit breakers

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Functions (3)

n Controln Observing parameters

• Current, voltages, temperatures, status, …n Information are transmitted to the Ground by telemetry for mid-term and long-

term monitoring n Information are transmitted to the On-Board Computer for real-time monitoring

n Commandn Configuration setting (nominal, safety, recovery, …)n Parametersn ON/OFF

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System drivers (1) / Synthesis

The orbitn Low Earth Orbit (LEO), geostationary (GEO), Mean Earth Orbit (MEO), Sun

Synchronous Orbit (SSO), Sun Centric (Interplanetary), …

The missionn (Life) durationn Energy budget

n Mission profilesn Payload needsn Max and Mean powern Orientation (attitude) of the satellite

n Reliability requirements

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System drivers (2) / Orbitsn Eclipses

Ø Sun = 109 x Ø Earth

n Geo • Less then 1% of mission duration• Only during equinoctial periods• From few to 72 min max

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System drivers (3) / Orbitsn Eclipses

n Leo• High variability versus the orbit selection• Up to 40 % of eclipse duration• Thousands of cycles along mission duration

0 200 400 600 800 1000 1200 1400 1600 1800 20000

5

10

15

20

25

30

35

40

Eclip

se D

urat

ion

(min

)E

clip

se D

urat

ion

(min

)

DaysDays0 0 35 min35 min

Period: 96.5 minPeriod: 96.5 min

0 200 400 600 800 1000 1200 1400 1600 1800 20000

5

10

15

20

25

30

35

40

Eclip

se D

urat

ion

(min

)E

clip

se D

urat

ion

(min

)

DaysDays0 0 35 min35 min

Period: 96.5 minPeriod: 96.5 min

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System drivers (4) / Orbitsn Radiation sources

n Trapped electronsVan Allen belts

n Trapped protonsVan Allen belts

n Sun protonsSun eruptions

n Space heavy ionsCosmic rays

n µ-meteorite

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System drivers (5) / Orbitsn Radiation sources

n Trapped electronsVan Allen belts

n Trapped protonsVan Allen belts

n Sun protonsSun eruptions

n Space heavy ionsCosmic rays

-> The radiation environment has a direct impact on the definition & sizing of EPS

EffectsEffects

Total doseTotal doseDecreasing of semi-conductor performances up to destruction

SA cells, Mosfets, Bi-polar transistors, …

S.E.E.S.E.E.Transient effect on semi-conductors,

may lead to its destructionMosfets, Memory, Amplifiers, …

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System drivers (6) / Orbitsn Solar flux, which decreases with the square of the distance to the sun

1.36 1031.001.0Earth

5820.531.5Mars

48.7115.2Jupiter

13.599.5Saturn

9.3 1030.380.39Mercury

1090Sun

Solar fluw (W/m2)

Radius (Earth)

Distance (AU)

0.860.1839.5Pluto

1.5430.1Neptune

3.6419.2Uranus

2.6 1030.950.72Venus

1 UA = 149 597 870 km1 UA = 149 597 870 km

Earth radius = 6378 kmEarth radius = 6378 km

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System drivers (7) / Orbitsn Solar flux, which greatly depends of the selected orbit

0 200 400 600 800 1000 1200 1400 1600 18000

500

1000

1500

days

Ava

ilabl

e P

ower

(W/m

²)

0 200 400 600 800 1000 1200 1400 1600 18000

200

400

600

800

1000

1200

1400

days

Ava

ilabl

e P

ower

(W/m

²)

SA flux (sun-synchronous orbit)- > Min SA flux = 1220 W

SA flux (polar orbit)- > Min SA flux = 520 W

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System drivers (8) / Missionsn (Life) duration

n From few minutes (launchers) to 15 years (Geo)

n Ageing drifts shall be assessed on each EPS constituent / Even some manufacturers may not be qualified for long term missions (e.g. ABSL batteries)

n Impact on total radiation dose & nb of thermal cycles

n Reliability requirementsn EPS may be requested to be

• SPF free• No single failure may lead to the loss of mission• Note: for human mission, no combination of two failures may lead to the loss of

mission• Not reliable

• E.g.: in µSAT, any failure may lead to the loss of mission-> Important impact on system architecture (definition of redundancy) and on system cost

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System drivers (9) / Missions

n Energy budgetn Mission profilesn Payload needs

• TV broadcasting points a zone of the Earth• Science satellites may point any zone of the sky• Military satellites may point any zone of the earth and shall be very agile

n Max and Mean power (in sunlight and in eclipse)n Orientation (attitude) of the satellite. The attitude constraints directly drive the

sizing of the primary and secondary sources: impacts on• Eclipse duration• SA flux• Payload power available (in sunlight and in eclipse)• Definition of recovery / safety attitudes of the S/C• …

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EPS equipments glossary

MPPTMaximum Peak Power Tracking

BCRBattery Charge Regulator

SSPCSolid State Power ControllerFCLFold-back Current Limiter

URBUnregulated BusLCLLatchning Current Limiter

State of Charge

Solar Array Drive Mechanism

Solar Array

Sequential Switching / Shunt Regulator

Regulaterd busPower Subsystem

Power Conditioning & Distribution Unit

Power Conditioning Unit

Equipment

RBDoDDepth Of DischargeS3REoCEnd of Charge

SAEoDEnd of Discharge

SADMEoLEnd Of Life

PCDUBoLBegin Of LifePCUBDRBattery Discharge Regulator

Equipment

SoCEPSElectrical Power Subsystem

PSSCVConverter

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EPS in practice (1)

TELECOM 2 (1991)Telecommunication – P = 3 kW

12 %132Battery NiH2

9 %100Solar Array

27 %27 %296296Power TOTALPower TOTAL

4 %43Power System (incl. SADM)

100 %1100Satellite dry mass

Satellite mass ratio

Mass (kg)

TELECOM 2

2 %21Distribution

Data from CNES

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EPS in practice (2)

JASON 1 (2001) Mini satelliteOceanographic Observation satellite – P = 500 W

10 %45Battery NiCd

9 %42Solar Array


2 %10Power System (incl. SADM)



Mass (kg)

Jason 1

6 %29Distribution

Data from CNES

Illustration CNES

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EPS in practice (3)

DEMETER (2004) Micro SatelliteScience (Geodesy) satellite – P = 110 W

4 %4Battery LiIon

6 %6.5Solar Array


6 %6.5Power System (incl. SADM)



Mass (kg)

DEMETER

Data from CNES

Illustration CNES

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Agenda





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Primary power sources / Solar cells & arrays

SA cellsn A solar cell is composed of a semi-

conductor material and converts photons to electrons

n Photovoltaic effectn The solar flux is reflected, absorbed

by the solar cell or crosses itn Every absorbed photon whose

energy is greater than semi-conductor gap is going to release an electron and to create a positive « hole » (lack of electron). This electron is part of the crystalline network

n Photons with excess energy dissipate it as heat in the cell, leading to reduced efficiency

n An electrical field is introduced in the cell in order to separate this pair of opposite charges

Illustration SPECTROLAB

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Semi-conductors properties (1)n Most common semi-conductors: silicium and arsenide-gallium (AsGa)

n Cubic crystalline structures

n Method of GaAs Growth: Metal Organic Vapor Phase Epitaxyn N-type contact (upper surface of the cell): multi-finger arrangement

• Efficient current collection• Good optical transparency• Connected at a bar along one edge of the cell

Silicium Arsenide Gallium

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Solar fluxn Semi-conductor gap is chosen to fit with space light wavelength

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Equivalent circuit diagram

n Each solar cell is equivalent to n a current source in parallel with n a capacitor (variable) and n a diode

VI

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Optimal working point – at max. available power

Psa characteristics

020406080

100120140160180200220240

0 4 8 12 16 20 24 28 32 36 40 44 48 52 56 60 64 68 72 76 80 84

U sa (V)

P sa

(W)

Psa-BoL-HotPsa-EoL-HotPsa-BoL-Cold

U-I characteristic

0.0

1.0

2.0

3.0

4.0

5.0

6.0

0 4 8 12 16 20 24 28 32 36 40 44 48 52 56 60 64 68 72 76 80 84

U sa (V)

I sa

(A)

Isa-BoL-HotIsa-EoL-HotIsa-BoL-Cold

Isc

Voc

Pmax

Pmax

Pmax is largely depending of temperature & ageing

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Efficiency


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Efficiency degradation factors

n Mission lifetimen Loss of power: 1% to 2% every day (depends of the orbit)

n Radiation effects

n UV n Meteorite impactn ATOX density

n Aggressive and corrosive environment (tied to the LEO) on cover glass protection and on exposed interconnection (oxidation of silver and then increase of resistivity)


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Concentrators / Advantages

n Concentrate SA flux on SA cellsn Reduce SA cells surfacen Based on reflectors or on lens

Direct Illumina tion

Reflected Illumina tion

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Concentrators / Drawbacksn Not compatible with large off-pointing angle

n Oblique rays can hit the reflectors two times and then they may be reflected back to space

n Off-pointing property of SA concentrator is not compliant with un-stabilized S/Cn Induces higher thermal constraints on cells and SA panel

n Outgassing may become critical

50

60

70

80

90

100

110

120

-0.08 -0.06 -0.04 -0.02 0 0.02 0.04 0.06-0.06

-0.04

-0.02

0

0.02

0.04

0.06

0.08Tempera ture map

Concentration at large off-pointing

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

0 10 20 30 40 50 60 70

Off-Pointing (arcdeg)

Concentration

Trough C=1.8Trough C=2Trunc Trough C=1.85CPC 10degCPC 24deg

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Solar arrays (1)

n A solar cell produces some hundreds of milliwatts

n A solar Array (SA) is composed of thousands cells assembled in series and in paralleln The network = cells +

interconnections + cabling + diodes

n A string = assembling of cells in series to obtain the desired voltage

n A section = strings in parallel to obtain the desired current

n Sections are independentIllustration Thales Alenia Space

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Solar arrays (2)

Protection against shadowing

Illustration Thales Alenia Space


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Solar arrays (3) – Types

n Fixedn Solar cells are glued on the

structure of the satelliten The power is limited by the

surface of the satelliten Deployable (fixed)

n Solar cells are glued on flaps (folded at launch and deployed in orbit)

n Difficult to manage the attitude constraints

n Deployable and mobilen 1-degree of freedom

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Solar arrays (3) – Types

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Panel, glue, coverglass, …n Substrate

n Kapton with glass – or carbon- reinforcement

n Glue, Adhesive n Fix SA cell on SA paneln Fix the coverglass on the celln Ensure electrical & thermal conductivity

n Panel (honeycomb)n Support SA cellsn Transfer heat to bottom siden Face high thermal gradientn Be compatible with deployment and orientation mechanisms

n Coverglass n Protect SA cell against ATOXn Protect SA cell against radiationn Limit the UV flux to the adhesive layer and to the cell by allowing suitable wavelength

selection, via a good optical coupling (between free-space and glass & between glass and adhesive)

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Solar Arrays (4) – mechanisms (hold-down & release)n SA are fold during launchn Deployment is

n Initiated by pyro actuation (or thermal knifes)n Controlled by the use of hinge mechanisms


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Solar Arrays (4) – mechanisms (orientation)n Mobile SA is controlled by SADMn Current is transferred to S/C main part via BAPTAn Tensioning wires – achieve minimum fundamental frequency of the array (AOCS

constraints)


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Solar arrays (5) – thermal interfaces

n SA cell temperature directly impacts its efficiency

n Temperature is linked ton The incoming flux

• Direct solar flux• Albedo• IR flux of the earthn The outcoming flux

• Flux reflected by the cells• Power delivered to the satellite• IR flux of the front and rear part of the SAn Sizing of Solar Arrays requires simulation tools

• e.g. the delivered power is function of the temperature, which is itself function of the power

n Temperature cycling (~100 °C in few minutes) induces thermal stress (differential expansion between substrate and cell, which are made of different material) on interconnections (major array failure hazard)

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Solar arrays (6) – performances

n Typical performances after 15 years in GEOn Silicium: 100 W / m2

n High efficiency silicium: 130 W / m2

n AsGa (mono junction): 170 W / m2

n AsGa double junction: 200 W / m2

n AsGa triple junction: 240 W / m2

n Power / kg:n Silicium or AsGa/Ge: 40-50 W/kgn Multi junctions: 50-60 W/kg

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Primary power sources / Fuel cells

Electromechanical devices performing a controlled chemical reaction (oxidation) to derive electrical energy (rather than heat energy)

n Advantagesn Minimal thermal changesn Compact and flexible solutionn Production of water (manned mission)

n Drawbacksn Need of fuels: hydrogen & oxygen

yielding water as the reaction product

n Used for Shuttle orbiter, lunar rover, …

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Typical current-voltage curve for a hydrogen/oxygen fuel cell

Performance summary of fuel cells for space use

550Goal (lightweight cell)> 40 000 h110Alkaline technology

Not drinking water240 h33Gemini

> 3000 h> 40000 h

2500 h

Operation

15 min start-up / instantaneous shutdown

Operated at 505 K24 h start-up / 17 h shutdown

Comment

367Alkaline technology110 – 146SPE technology

25Apollo

Specific power (W/kg)

System

275Shuttle

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Use of fuel cell as « secondary power source »n Regenerative fuel cells (100 kW system power) electrolyze of water is performed

during the ‘charge’ cycle thanks to primary source power

n Advantagen Lower SA power need thanks o judicious sizing of the fuel

n Drawbackn Lower efficiency (50 – 60 %) than battery

n Interesting for LEO operations where atmospheric drag is important (very low orbits) -> reduction of propellant used for orbit control

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Primary power sources / RTG

Deep-space missions (further than Mars) or Military usen Long time missions, not-compatible with fuel cellsn Far from Sun, not-compatible with SA

n Decrease of SA flux partially compensated by increased of cell efficiency due to decrease of temperature (rE/rSC)1.5

-> Use of radioactive decay process, use of thermoelectric effect

Thermoelectric effectn Generation of a voltage between (semi-conductor) materials maintaining a

temperature difference. Power function of:n Absolute t° of hot junctionn T° difference between materialsn Properties of materials

n Low efficiency (< 10 %) -> removing waste heat may be an issue

n Heat source: spontaneous decay of a radioactive material, emitting high-energy particles, heating absorbing materials

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Primary power sources / RTG

Advantagesn Power production independent of S/C orientation &shadowingn Independence of distance from Sunn Low power level may be provided for long time periodn Not susceptible to radiation damagen Compatible with long eclipse (e.g. lunar landers)

Drawbacksn Affect the radiation environment of S/C

(deployment away from the main sateliite bus)

n Radioctive source induce safety precautions in AIT

n High t° operation required -> impact thermal environment of S/C

n Interfere with plasma diagnostic equipment (scientific missions)

n Environmental risk in case of launch failure or S/C crash

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Primary power sources / RTG & others

Example of RTG

n Cassini (Saturn mission) 628 W 195 W/kgn Galileo probe/Ulysses 285 W 195 W/kgn Nimbus/Viking/Pionner 35 W 457 W/kgn Appolo lander 25 W 490 W/kg

73 W 261 W/kg

Nuclear fissionn Use of nuclear fission process (as for terrestrial nuclear power plants)n Fissible material (e.g. uranium-235) used to drive thermoelectric converter as RTG

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Primary power sources / Others

Solar heat

n Use of Sun energy to drive a heat engine and then a rotary converter to electricity or a thermoelectric converter

n Concept interesting for space stationn Reduced drag (reducing area of SA panels)n Reduced maintenance effort

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Agenda





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Secondary power sources

Accumulatorsn Electromechanical devices performing a controlled chemical reaction to derive

electrical energyn Critical parameters

n Charge/discharge raten Depth of Dischargen Extent of over-dischargingn Thermal sensitivity to each of

these parameters

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Comparison of performances

907575Energy efficiency4.01.61.55Max. voltage (A)2.71.01.0Min. voltage (A)

1.5,2.2,26,4030 ⇒ 3504 ⇒ 50Capacity (Ah)

⇒ C⇒ C⇒ 2CDischarge current (A)

100-13055-6530-40Energy/kg (Wh/kg)

5 years at 40 % of DoD


⇒ C/8 (GEO) ⇒ 0.7C (LEO)

[0;+10]

1.25

80

NiH2



C/10 ⇒ C/3

[+15;+25]

3.5

200-250

Li-Ion

⇒ C/10 (GEO) ⇒ C/2 (LEO)

Charge current (A)


Life duration in Geo

[-5;+15]Working temperature (°C)


Life duration in Leo

110Energy/l (Wh/l)

NiCdType

1.25Discharge voltage mean (V)

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Batteryn Role: support the Solar Array during

n LEOP phasesn Eclipsesn Loss of sun pointingn Peak power demandsn …

n Series / parallel assembling of accumulator cellsn In series to reach the desired voltage

• 22-37 V in LEO• EUROSTAR 2000: 42.5 V• EUROSTAR 3000 & SPACEBUS 3000: 50 V• SPACEBUS 4000: 100 V

n In parallel to reach the desired capacity

Illustration SAFTIllustration SAFT

Illustration SAFT

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Battery failure modes

n Open circuitn Loss of battery

n Short failuren Degradation of the voltage

YesYesNiH2

YesNever metNiCd

Short failureOpen circuit

Yes ? Not very well known

Yes ? Not very well known

Li-Ion

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Battery protections

n By-passn To isolate the failed cell (open or short mode)

n Balancing (Li-Ion)n Voltage balancing at cells level to improve battery end-of-charge voltage and

increase cell life-timen Counterbalancing of cells mismatching

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Battery accommodation

n Mechanical & thermal driversn The structure shall be rigid (launch environment)n The structure shall allow the homogeneity of the temperature

ZONE 2b - LOWER RADIAL PANEL

1

10

100

1000

10000

1 10 100 1000 10000Frequency [Hz]

Acc

eler

atio

n [g

]

Lower Radial Panel - Qualification (Flight+3dB) [g]

Lower Radial Panel - Flight [g]

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Battery sizing

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Sampling of accumulators

70 %

57 Wh/kg

2108 g

120 Wh

1.36 V

89 Ah

SAFT NiH293 AN

90 %

126 Wh/kg

1107 g

140 Wh

3.6 V

38.6 Ah

SAFT LiIonVOS140

90 %

141 Wh/kg

155 g

22 Wh

3.6 V

6.1 Ah

SAFT LiIonMP76065

41.2 g1610 gMass

126 Wh/kg34 Wh/kgEnergy/kg

5.2 Wh55 WhEnergy

90 %70 %Efficiency

1.4 Ah46 AhCapacity

SONY LiIOn18650HC

SAFT NiCdVOS 40

BoL

3.7 V1.2 VMean voltage

Data CNES

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Sampling of batteries

910X520X300mm490X380X290mm863X441X310mm467X261X260mmDimensions

12s-4p VOS140

11s-2p VOS140

27s-1p 93AN24s-1p VOS40Configuration

29 Wh/l

52 Wh/kg

66 kg

3415 Wh

37 V

93 Ah

EURASIASAT

59 Wh/l

93 Wh/kg

34 kg

3168 Wh

39.6 V

80 Ah

STENTOR

47 Wh/l

92 Wh/kg

72 kg

6620 Wh

43 V

154 Ah

SKYBRIDGE

47.4 kgMass

28 Wh/kgSpecific energy

1325 WhEnergy

42 Wh/lDensity

46 AhCapacity

SPOT5BoL

29 VMean voltage

Data CNES

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Sampling of batteries

4 kg40,4 kgMass

226 x 166 x 95 mm2 x (355 x 295 x 180 mm)Dimensions

105 Wh/kg105 Wh/kgSpecific energy

420 Wh4200 WhEnergy

118 Wh/l170 Wh/lDensity

14 Ah140 AhCapacity

8s-10p 18650HC8s-100p 18650HCConfiguration

µSATPLEIADESBoL

30 V30 VMean voltage

Data CNES

Illustration SAFT

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Agenda





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Power management, control & distribution / Architecture

Block diagram

-> Operation with primary & secondary power sources whose characteristics are changing with time and conditions of operations

PCDU

SAconditioning

SA

BAT

BCR /BDR /

Switches

Distribution User's

BusDivided in

sections

1 or 2 batteries

Monitorings

DET, MPPT, …

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Bus voltage selection

n Many standardsn 28V, 50 V, 65 V, 100 V, …n Even Ac busses are used for high power spacecrafts (e.g. ISS)

n Choice is based onn Bus power

• Recommended ESA rule: P < U2/0.5 for bus impedance reasons• High bus voltage means

• Less current• Simplification of harness• « High » voltage management at equipment level (SA, battery, PCDU, …)

n Payload flight heritage

-> Some architecture may even requires two buses !!

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Conditioning architecturen Regulated bus

n Voltage variation is limited to about +/- 1 V whatever the satellite modes

n Need of dedicated electronics to manage the battery discharge• Substantial power dissipation inside the PCDU during

eclipse phasen Unregulated bus

n Bus voltage is imposed by the battery voltage• Impact on all DC/DC converters efficiency

n Semi-regulated busn Regulated bus in sunlight only

n Choice is based onn User’s need (mission)

• Scientific payloads may require regulated bus to fulfill their precisions

• Thermal stability of some specific loads may requires regulated bus (thermal management is easier in that architecture)

n User’s flight heritage

PCDU

unregulated 28V

power distribution to users

BAPTA

Li-Ion Battery

MPPTcell PDU

regulated 28V

power distribution to users

BAPTA

Li-Ion Battery

S3RS4R

BDR

PDU

PCDU

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Conditioning topologyThe conditioning concerns the way the power coming from the solar array is used to be

delivered to the different users of the spacecraft, as well as to the battery in order to guarantee the battery recharge

n Direct Energy Transfer (DET)n DET operates at the bus voltage and extracts the available power from the solar

array for this precise voltage• Simplest solution

n PWM control of SAn DC/DC converter implemented between SA and bus

n Maximum Power Point Tracking (MPPT)n MPPT can operate in a wide range of voltages to track the maximum available

power from the solar array, converts the (VMP, IMP) into (Vbus, Ibus) and is particularly interesting in case of sensitive flux variations• More complex and dissipative solution

n Choice is based onn Mission & orbit

• SA flux variation (agility of the S/C, interplanetary missions, …)• SA temperature variabilityn …

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Battery managementn Architecture

n Centralized• Performed by OBC• PCDU functions limited to monitoring

n De-centralized at PCDU level• Autonomous

• PCDU ensures battery charge & protections in a reliable way• Voltage tapering• Protection against over-charge, over-discharge, over-temperature, …

• Partially autonomous• PCDU ensures battery charge & protections• OBC is responsible of PCU re-configuration in case of internal failure

n Any intermediate solutions between these two extremesn Choice based on

• Price• Satellite reliability need• …

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Battery managementn Functions

n Voltage / current control during chargen Current limitation during dischargen Monitoringn Protections

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Distribution architectureDistribution concerns the way the power is distributed from primary & secondary sources

to user’s through PCDU. To avoid failure propagation in case of user’s short failure, these lines shall be protected by

n Fusen Simplest solutionn Imposes all the user’s to be compatible with bus transients induces by fuse

blowingn Imposes the need of extraction during AIT phase

n SSPCn Flexible solutionn ON/OFF switching capabilityn Control of fault current

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Distribution architecture / some definitions

n LCLn Latching Current Limitern Limits current at user’s switch ON or short

failure during limitation timen Trips-OFF if limitation time is exceededn ON/OFF command capability

n FCLn Fold-back current limitern Essential load (e.g. OBC)n Limits current at user’s switch ON and

during short failure (with decreasing level)n PO-LCL

n Permanent-ON LCLn Essential load (e.g. OBC)n LCL + automatic periodic re-arming

n …

dI/dt < 1 A/µs

Inom

Ilim.minIlim.nomIlim.max

Imax< 1.2 x Ilim.max

Reaction time< 1µs

Trip-off time

I

t

Current fallsdepends on

loadscharacteristics

(L / C) only

Voltage

Current

FCL I - V CHARACTERISTICS

Ilimit - max Ilimit - min

Vbus

Inom

Iclass Voltage

Current

FCL I - V CHARACTERISTICS

Ilimit - max Ilimit - min

Vbus

Inom

Iclass

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Other constituents of PCDU

n Command of mechanismsn SADM motor drivern Antenna motor drivern …

n Command on deployment n Actuation of pyron Actuation of thermal knifes

n Li-Ion battery cells managementn Acquisition of thermistorsn …

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Battery management / Li-Ion battery cells management

n Cells balancingn Voltage balancing at cells level via the actuation of a shunt in parallel with the

cell to slightly discharge itn Cells by-pass

n Actuation of electro-mechanical device allowing to short circuit a failed cell and avoid failure propagation at battery level (arm / fire circuitry)

n May be integrated at battery level or at avionic level (PCDU or not)n Some Li-Ion batteries do not need cells balancing thanks to battery inner property

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Power management, control & distribution / PC(D)U

Examples µSAT

n Low power: 260 W / Low voltage : unregulated bus (22-37 V)

n Solar Array regulator: Boost convertern Not reliablen Distribution functions

n LCL, Pyron DC/DC for secondary (+5, +-15,+20

V) + LCL protectionn Adaptability of the distribution by

paralleling of LCL’sn CNES/Astrium/TAS-F Myriade platform

baseline

Illustration CNES

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Examples Scientific & earth observation

n Large flexibility neededn Modular structuren Large flexibilityn Redundancy (tolerant to one failure)n Bus Power : 500 W to 4200 Wn Bus voltage : up to 50 V, non-regulated or regulatedn Solar Array Regulation : MPPT or DET (S3R or S2R)n Lithium Cells Management : cells voltage balancing and by-pass electronicsn Distribution : LCLs, FCLs, Relays+Fuses, Heater Switches, Pyro Electronicsn TMTC : MIL-1553B bus or other

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Examples Scientific & earth observation

Illustration CNES

Illustration ESA

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Examples interplanetary cruise (-> Sun)

n High power PCDU (up to several kW)n To accommodate with electrical propulsion

n Bus voltage: up to 100 Vn Solar array regulation: MPPT

n To accommodate large SA flux variationn Low mass & volume requirementsn Redundancy (tolerant to one failure)n Distribution : LCLs, FCLs, Relays+Fuses, Heater Switches, Pyro Electronicsn TMTC : MIL-1553B bus or other

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Examples Geo low power

SPACEBUS 3000 PCUn Full regulated bus 5.5 kW / 50 V n Solar array regulation: S3Rn No distribution function (PCU only)

n Flight heritage : 28 PCU’s, 1 480 000 h

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Examples Geo high power

SPACEBUS 4000 PCUn Full regulated bus 6 to 21 kW / 100 V n Solar array regulation: S3Rn No distribution function (PCU only)

n Flight heritage : 9 PCU’s, 130 000 h

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Agenda





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Power budget

Study case

n Study of a micro satellite to target ship based and ground based radarsn Lifetime: 5 yearsn Orbit: Leo

n Payload requirementsn Acquisition in sun & eclipse phasesn Bus power of 626 W

• Max power to be considered• Sum of all user’s needs (AOCS, payloads, emitters, receivers, thermal control…)• Worst case consumption in all satellite phases (acquisition, data transmission, night&

day modes, seasons variation on thermal control, …)• Excluding LEOP phases• Excluding EPS needs

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Power budget

Orbit selectionn Altitude trade-off

n Lower than 1000 km (to avoid Van Allen belts impacts on radiation leveln Above 500 km to ensure that the cluster altitude can be maintained during

lifetime (atmospheric drag effect)n Instrument precision is better at low altitude but instrument coverage increases

with altitude-> Circular orbit of 600 km altitude has been selected among several candidates (out

of the scope of this study case, based essentially on payload needs)n Inclination trade -off

n Polar orbit for best possible coverage worldwiden Sun-synchronous orbit as other candidate

30 min

98 ° (sun-synchronous)

0.001 (circular orbit)5820 s600 km

90 ° (polar orbit)Inclination0.001 (circular orbit)Eccentricity

21.3 minEclipse duration

600 kmAverage heightOrbit characteristics

5820 sPeriod

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Power budget

Orbit selection / Inclination trade-off

0 200 400 600 800 1000 1200 1400 1600 18000

500

1000

1500

days

Ava

ilabl

e P

ower

(W/m

²)

0 200 400 600 800 1000 1200 1400 1600 18000

200

400

600

800

1000

1200

1400

days

Ava

ilabl

e P

ower

(W/m

²)

SA flux (sun-synchronous orbit)- > Min SA flux = 1220 W

SA flux (polar orbit)- > Min SA flux = 520 W

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Power budget

Orbit selection / Inclination trade-offn EPS sizing shall consider worst case conditions of illumination and EOL photovoltaic

efficiency of SA cells. This leads to the following data (worst case figures).

n Note that photovoltaic efficiency EOL/BOL ratio takes into account the following elements (SA panel manufacturer data)n 5-years mission lifetimen radiation effectsn UV and meteoritic impactn effect of ATOX density (aggressive and corrosive environment tied to the LEO)

on cover glass protection

111

520Polar

76.5 %EOL/BOL ratio28 %BOL SA cell efficiency

261Total available SA power (W / m2)

Sun-synchronous1220Minimum SA flux (W/m2)

Cell manufacturer data

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Power budget

Batteryn Additional hypotheses considered for EPS trade-offs

n Battery recharge duration = 90 % of sunlit durationn Battery dissipation (at battery level)

• 25 W (discharge)• 15 W (charge)

n PCU to battery harness losses = 3 %n Max DOD of 40 % considered following

• Orbit characteristics (period and eclipse)• Mission duration

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Power budget

EPS sizing / bus voltage trade-offn 28 V

n Compatible with bus power (< 1 kW)n High hardware heritage

n 50 Vn Reduced current levelsn Reduced harness & power dissipations

EPS sizing / bus regulation trade-offn Regulated power bus – main hypothesis

n Unregulated power bus – main hypothesis

Note: PCDU low level consumption: 30 W for both configurations

S3R (SA => main bus) conversion efficiency 98 % S4R (SA => battery) conversion efficiency 95 % BDR (battery => main bus) conversion efficiency 92 %

MPPT (SA => main bus and SA => battery) conversion efficiency

95 %

Battery => main bus losses (harness) 1 %

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Power budget

EPS sizing / bus regulation trade-off (sun-synchronous orbit)

Slight advantage for URB coupled with lower PCDU mass. If no specific requirement on payload (including EMC), URB is selected.

Regulated bus Unregulated bus

Users’ power in eclipse (W) 626 PCDU losses during eclipse (W)1 84 36 Satellite power requirement in eclipse (W) 710 662 Eclipse duration (min) 30 Discharge dissipation losses (W) 25 Battery recharge duration (h) 1.005 Recharge dissipation losses (W) 15 Battery recharge power requirement (W) 381 357 Harness losses (W) 11 11 S4R / MPPT losses (W) 20 18 SA power requirement for battery recharge (W) with 10 % margin 453 424

Battery useful cycled energy requirement (Wh) 355 331

Maximum acceptable DOD 40 % Discharge dissipation losses ( %) 2.5 % BOL battery energy requirement (Wh) @ 40 % DOD with 10 % margin 1001 933

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Power budget

EPS sizing / Conditioning topology trade-offn MPPT

n Conversion efficiency 95 %n Ability to track the maximum power

whatever the battery state is (charged, discharged, with or without failure, …).

n DETn Conversion efficiency 97.5 %n Since DET extracts SA power at fixed

battery voltage [22 V; 37 V], SA electrical efficiency is never perfectly optimized, leading to a mean value 10 % lower than maximum value (typical value, function of SA sizing & battery failure modes)

Note: Considering 28 V URB with 40 % DoD, battery voltage is comprised between 22 & 37 V

Psa characteristics

020406080

100120140160180200220240

0 4 8 12 16 20 24 28 32 36 40 44 48 52 56 60 64 68 72 76 80 84

U sa (V)

P sa

(W)

Psa-BoL-Hot

Psa-EoL-Hot

Psa-BoL-Cold

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Power budget

EPS sizing / Conditioning topology trade-offn MPPT

MPPT offers n About 9 % SA surface reduction (compared with DET solution)n Greater flexibility towards satellite FDIR

butn imposes some constraints on thermal interfaces (dissipation inside PCDU is higher)n Impacts PCDU definition (more complex solution)

MPPT DET SA efficiency (W/m2) 261 235 Users’ power in sunlit (W) 626 PCDU losses during sunlit (W)1 63 46 Satellite power requirement in sunlit (W) 689 672 Battery recharge power requirement (W) 386 386 Total SA requirement (W) with 10 % margin 1182 1163 Minimum SA surface requirement (m2) 4.5 4.9

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Agenda

1. Introductionn Presentation of Thales Alenia Spacen Presentation of myselfn EPS – general information




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Conclusions

The design of any Power Subsystem is strongly linked with Systemanalyses (Attitude & Orbit, Mission, Operations)

The electrical architecture of spacecrafts is not standardn Unregulated or regulated Power busn Voltage (28 V, 50 V, 100 V, …)n Conditioning (S3R, MPPT, …) n Protections (reliable or not)n Distribution (fuse, LCL, …)n …

and shall be adapted nearly on case by case ….

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Annex

Thales Alenia Space ETCA – Proposition de stages / TFE

n Senseur de courant (large bande, filtrage numérique)n Correcteur numérique stochastiquen Onduleur multi-leveln Cellule de commutation de puissance résonante (série/shunt, CALC)n Circuit de commande de grille (isolation, résonance, high dV/dt immunity)n Convertisseur à étalement de spectren Convertisseur DC/DC de type SEPICn Topologie DC/DC utilsant les CPI (Composants Passifs Intégrés)n Blindage des noyaux magnétiquesn Dispositif anti-saturation pour transformateurn Topologies de filtres EMC actifs/passifsn Transformateurs ‘Coreless’ (puissance / bas-niveau)n Dispositif de détection de température fiable « Thermal lockout »n Elimination de « RHP zero » par modulation du flanc montant

105Ref : VLR/080066Date : 28.01.08

10/04/2007

Thales Alenia Space ETCAn Tél. : + 32/71 44 22 11 n Fax : + 32/71 44 22 00n www.thalesaleniaspace.comn E-mail : [email protected]

Vincent Lempereurn PCU/PCDU product managern Tél: +32/71 44 28 89n E-mail: [email protected]

http://www.thalesaleniaspace.com

mailto:[email protected]

mailto:[email protected]

electrical power systems - uliege · 12 ref: vlr/080066 date : 28.01.08 introduction / tas-etca...

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