Electrical Power Systems

Electrical Power Systems

Vincent Lempereur Ref :

VLR/080066

Date :

21.10.13

Agenda 1. Introduction n Presentation of Thales Alenia Space Belgium (ETCA) n Presentation of myself n EPS – general information

2. Primary power sources n n n n

Solar cells & solar arrays Fuel cells RTG Others

3. Secondary power sources - batteries 4. Power Management, Control & Distribution n Architecture n PCU / PCDU

5. Power budget – practical exercise 6. Conclusions

2 Ref :

VLR/080066

Date :

21.10.13

Agenda 1. Introduction n Presentation of Thales Alenia Space Belgium (ETCA) n Presentation of myself n EPS – general information





3 Ref :

VLR/080066

Date :

21.10.13

THALES

A global company with 67.000 employees and 14,2 billion in revenues in 2012 Aerospace and Transport Defense and Security

45%

55%

Aerospace & Transport

Avionic Systems

In-Flight Entertainment

Top Deck Avionics Air Traffic Control Telecommunication Urban Transport Systems Satellites Signalling Systems

Simulators

4 Ref : VLR/080066 SOC-PPT-110088-05-00-TAS ETCA UK 15.03.2013 Date : 21.10.13

Air Traffic Earth Observation Main Line Rail Surveillance Radars Satellites Signalling Systems

Défense & Sécurité

Watchkeeper

ACCs LOC 1

SAMP-T

Syracuse III

Hypervisor

Damocles

Ground Master 400

Sophie MF

Flexnet SDR

Sonar 2087

Thales Alenia Space : A world player in each of the space market A global offer from equipment to end-to-end space systems (Design, manufacturing and delivery) Telecommunications

Fixed / Mobile Broadband Dual / Military Secured

O3B, Palapa-D1, RascomStarQaf 1R, Satcom BW 2a/2b, Sicral 1B, Sicral 2, Syracuse, 1/2/3, Thor 6, Turksat 3A, W2A, W3B, W3C, W3D, TKM, 8WB, …


Observation

Climat change Meteorology Oceanography Intelligence Surveillance

Meteosat/MSG/MTG Jason, Calipso, COSMO-SkyMed, Sentinel 1&3, Envisat, ERS, GOCE, Helios, IASI, Pleiades, SPOT, CSO, Gokturk, …

Navigation

Exploration / Science

Localization Aeronautical Communications Data collect

Planetology Fundamental physics Astronomy Human spaceflights Space transportation systems

EGNOS, Galileo, MTSat, …

Herschel,Planck, Exomars, Alma, Corot, Cassini-Huygens, Node 2&3, Columbus, Cupola, MPLM, Automated Transfer Vehicle (ATV), Ariane 5…

A unique combination of expertise covering the full value chain A joint-venture between two leaders Thales: 67% Finmeccanica: 33%

Equipments


Payloads

Finmeccanica: 67% Thales: 33%

Satellites

Systems

Services

Key figures : 2012

7200 employees 10 sites in 5 countries 2 Bn€ sales in 2012

A strong industrial presence in Europe Commercial representation in multiple countries Certifications : ISO 9001, EN9100, AQAP 2110, CMMI3, ISO 14001


Customers all over the world

Boeing Comdev EMS Globalstar ILS Intelsat JPL Lockheed Martin Loral Space & Communications MDA NASA Orbital SES Americom Sirius Radio Space Systems Loral Northtrop Grumman XM Satellite Radio Worldspace


Arianespac e ASI Bosch CNES French MoD DLR EADS ESA

Nahuelsa t Star One

Eumetsat Eurasiasat Europe*Star European Union Eutelsat France Telecom Fire Brigades

Asecna Celtel Divona Telecom Globacom

Galileo Globecast Italian MoD Hispasat Protezione Civile Safran Satlynx

Libertis Gabon Rascom Telecom Vodacom

SES Astra SES Global SES Sirius SNCF Telenor Turksat Trenitalia Vigili del Fuoco

Arabsat ELOP IAI Noviasat TDCA Yahsat

ISR O

Gascom Krunitchev NPO/PM RSCC RKK Energia Mitsubishi NEC/Toshib a Sharp APT ADD CASC/CAST Chinasat ETRI Indosat KT NFA ShinSat Sinosat TDCA Toptry

Thales Alenia Space Belgium (ETCA)

A WORLD LEADER IN POWER ELECTRONICS FOR SATELLITE AND LAUNCHERS

50 years of experience


TAS Belgium (ETCA) – Key figures

Turnover (2012) : n 86 M€ (100% for export) n 25% launchers n 75 % satellites

More than 500 people with various expertises and profils 45 % engineers and managers 22 % workers 33 % tehnicians and administratives


Satellites : PCU / PCDU Worldwide Leader in electrical power conditioning and distribution n A portfolio adapted to any kind of satellites : n n n

Telecom, navigation, observation, scientific, … Any orbit (LEO / MEO / GEO) Power range : from 250 W to 21 kW

n Worldwide Customers n n

China (DFH4), Russia (Yamal), Argentina (ARSAT), Israel (AMOS), Japan (NT Space), India (ISRO), … Exclusive supplier of Thales Alenia Space for Spacebus satellites and constellations (Globalstar, O3B, Iridium Next)

n Main supplier of ESA and CNES for scientific and observation satellites n


Galileo, XMM, Herschel, Planck, ATV, Sentinelle 1 & 3, Myriade, …

Satellites : EPC / LC-TWTA

Key player in Electrical Power Conditioner (EPC) for Travelling Wave Tube Amplifiers (ATOP/LCTWTA) "Single" version More the 800 FMs ordered Nearly 35.000.000 hours of flight heritage Express AM/44, Herschel Planck, Yahsat, Loutch, Nilesat, Palapa, Sentinelle 1/2/3, Sicral 2, CSO, W3B/C/D, Exomars, MTG, …

"Dual" version Qualified at end of 2012 Innovative functionalities : Flexible LC-TWTA (worldwide exclusivity) TKM, Eutelsat 8WB


Satellites : PPU

European Leader in Power Supply for electrical propulsion satellite thrusters n Compatible to european satellites n

41 FMs ordered for Astrium (EUR3000), Thales Alenia Space (SB4000), OHB (SGeo), IAI (AMOS)

n Compatible with european thrusters (SNECMA PPS1350) and russian thrusters (Fakel SPT100) n Large increase foreseen on this market n

New generation (MK2) under qualification Increased power capability (up to 2.5 kW)

n

Future generation (MK3) under development For "full electrical" satellites (up to 5 kW)

•


•

Satellites : other products Avionics n Motor Drive Electronics (for solar arrays, antenna, thrusters, , ...) n Avionics (power distribution, pyro-commands, heaters, battery management, ...) n Specific Power Supply (Pulsed Power Supply for radars and altimeters) n DC/DC Converters


Launchers : Ariane 5

Main supplier for Ariane 5 n Development and production of the control and safety units of the launcher n 25 equipments and modules per launcher (exclusive supplier) n More than 50% of Ariane 5 electronics n 6 to 7 launchers per year


Launchers : Soyuz (launched from Kourou)

Exclusive supplier of the safeguard system of Soyuz (launched from Kourou) n 1 complete system per launcher n 2 to 4 launchers per year since 2011 n The only one non russian embedded system of the launcher


Launchers : Ground Equipments

Main supplier of control benches for Ariane Supplier since Ariane 1 (1970’) Control of the electrical function and the fluid during integration, launch chronology, management of the fluids of the launch pad.


A complete value chain

The A to Z of the products and services : from conception to delivery 280 engineers 42 000 m² of which 22 000 m² dedicated to manufacturing and tests 6 workshops of which 4 clean rooms (10 000 m²)


Some references Telecom Satellites : AMC12, AMC13, Apstar, Syracuse, Star One, HotBird 7, Chinasat, Express, Turksat, AM11-22-33-44, Koreasat,Ciel-2, W2A, W7, Satcom, BW, Sicral 2, Globalstar, Nilesat, Loutch, Palapa-D, Apstar 7B, Rascom, O3b, Arsat, Iridium, W6A, Artemis…

Observation, Scientific and Navigation Satellites: Meteosat, Smart 1, Pléiades-HR, Spirale, Stereo, Galileo, Herschel, Planck, Giove-B, Myriad, Sentinelle 1, 2, 3, Spot, Helios, Iso, XMM, Integral, Huygens, Soho,...

Launchers and Space Transports: Ariane 5, Soyouz, Vega, ATV, Expert, Colombus, ARD

Certifications : ISO 9001, EN 9100, AQAP 2110, CMM3 and ISO 14001 19 Ref : VLR/080066 SOC-PPT-110088-05-00-TAS ETCA UK 15.03.2013 Date : 21.10.13

Introduction / V. Lempereur Training n Electrical engineer from ULG (1996) n MBA in IEFSI (Lille)

Professional career in TAS-Belgium n Designer on PCU & PPU projects n SB4000 SBR & PCU, SB4000 PPU & FU n 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, MIRI n µSAT PCDU n PCDU/PCU product manager

20 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information A satellite is made of… P/F (Platform) n n n n

Mechanical & thermal structure Electrical system, avionic, propulsion On-board computer, software, remote control Energy sources: solar, batteries, fuel

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

21 Ref :

VLR/080066

Date :

21.10.13

P/F P/L

Introduction / EPS general information 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

22 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information General functional block diagram

Primary sources Power generation

Power management Regulation & Conditioning

Secondary sources Energy storage

23 Ref :

VLR/080066

Date :

21.10.13

Power distribution Distribution & protection

User’s Spacecrafts loads

Command & Control Interfaces

Introduction / EPS general information Functions (1) n Power generation n 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 storage n The energy is generally stored under a electro-mechanical form and retrieved under an electrical form n The storage of the energy is done by a secondary source, when the primary system’s energy is not available or insufficient Primary sources Power generation

24 Ref :

VLR/080066

Date :

21.10.13

Power management Regulation & Continionning



User’s Spacecarfts loads


Introduction / EPS general information Functions (2) n Conditioning and regulation n 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

n Distribution n To distribute the conditioned power to users n DC/DC voltage converters Primary Power management sources n ON/OFF switches Regulation & Power Continionning generation n Does not include the harness


25 Ref :

VLR/080066

Date :

21.10.13




Introduction / EPS general information Functions (3) Primary sources Power generation

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

• •

Fuses Circuit breakers

Power management Regulation & Continionning





n Control n Observing parameters

•

Current, voltages, temperatures, status, …

n

Information are transmitted to the Ground by telemetry for mid-term and longterm monitoring n Information are transmitted to the On-Board Computer for real-time monitoring n Command n Configuration setting (nominal, safety, recovery, …) n Parameters n ON/OFF 26 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Synthesis The orbit n Low Earth Orbit (LEO), geostationary (GEO), Mean Earth Orbit (MEO), Sun Synchronous Orbit (SSO), Sun Centric (Interplanetary), …

The mission n (Life) duration n Energy budget n Mission profiles n Payload needs n Max and Mean power n Orientation (attitude) of the satellite n Reliability requirements

27 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Orbits LEO (Low Earth Orbit) Scientific applications – Earth observation n Orbit n n n n

Ø Sun = 109 x Ø Earth Type: Circular Altitude: between 350 and 1000 km Duration:~2 hours Low sensitivity to radiations

n Eclipses n High variability versus the orbit selection n Up to 40 % of eclipse duration n Thousands of cycles along mission duration n Mission duration n 3 to 5 years n Example(s): Sentinel, CryoSat, …

28 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Orbits GEO (Geostationary Orbit): Telecom applications n Orbit Type: Circular n Altitude: 35786 km n Duration: 24 hours n Medium sensitivity to radiations n Eclipses n Less then 1% of mission duration n Only during equinoctial periods n From few to 72 min max n

n Mission duration n 15 years n Example(s): Spacebus based satellites, …

29 Ref :

VLR/080066

Date :

21.10.13

Earth

Sun

Introduction / EPS general information System drivers / Orbits MEO (Medium Earth Orbit): GPS applications n Orbit n n n n

Type: Circular Altitude: 1000 to 20000 km Duration: 12 hours Medium to high sensitivity to radiations (according to orbit height)

n Eclipses n Duration: 1 hour n Mission duration n 15 years n Example(s): GlobalStar, Galileo, Iridium, …

30 Ref :

VLR/080066

Date :

21.10.13

Earth

Sun

Introduction / EPS general information System drivers / Orbits Lagrange Point: Scientific applications - ESA n Points where the combined gravitational pull of two large masses precisely compensate the centripetal force required to rotate with them (analogy with the geostationary orbit) n Distance from earth for L1,L2: 1.5*106 km n Eclipses n None n Mission duration n 3 years n Example(s): Herschel (L2), Planck(L2), Gaia(L2),… Earth

31 Ref :

VLR/080066

Date :

21.10.13

Sun

Introduction / EPS general information System drivers / Orbits n µ-meteorite & debris

n

32

15 000 parts > 10cm n 300 000 parts < 10 cm n Large concentration between 700 & 1000 km New LOS (French rule) to avoid generation of new debris n Controlled desorbitation or n Parking in specific orbit with complete (propulsion and electronic) passivation (25 years in LEO, 100 years in GEO) Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Orbits n Radiation sources n Trapped electrons Van Allen belts

n

Trapped protons Van Allen belts

n

Sun protons Sun eruptions

n

Space heavy ions Cosmic rays

33 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Orbits n Radiation sources n Trapped electrons Van Allen belts

n

Trapped protons Van Allen belts

n

Effects Total dose Decreasing of semi-conductor performances up to destruction SA cells, Mosfets, Bi-polar transistors, …

Sun protons Sun eruptions

n

Space heavy ions Cosmic rays

S.E.E. Transient effect on semi-conductors, may lead to its destruction Mosfets, Memory, Amplifiers, …

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

34 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Orbits n Solar flux, which decreases with the square of the distance to the sun Distance (AU)

Radius (Earth)

0

109

Mercury

0.39

0.38

9.3 103

Venus

0.72

0.95

2.6 103

Earth

1.0

1.00

1.36 103

Mars

1.5

0.53

582

Jupiter

5.2

11

48.7

Saturn

9.5

9

13.5

Uranus

19.2

4

3.6

Neptune

30.1

4

1.5

Pluto

39.5

0.18

0.86

Sun

Solar fluw (W/m2)

Earth radius = 6378 km 35 Ref :

VLR/080066

Date :

21.10.13

1 UA = 149 597 870 km

Introduction / EPS general information System drivers / Orbits n Solar flux, which greatly depends of the selected orbit 1500 1400

1200

1000 Available Power (W/m²)

Available Power (W/m²)

1000

800

600

500 400

200

0

0

0

200

400

600

800

1000

1200

1400

1600

1800

days

0

200

400

600

800

1000

1200

1400

1600

days

SA flux (sun-synchronous orbit)

SA flux (polar orbit)

- > Min SA flux = 1220 W


36 Ref :

VLR/080066

Date :

21.10.13

1800

Introduction / EPS general information System drivers / Missions n (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 requirements n 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 37 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information System drivers / Missions n Energy budget n Mission profiles n 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

• • • • •

38 Ref :

VLR/080066

Date :

21.10.13

Eclipse duration SA flux Payload power available (in sunlight and in eclipse) Definition of recovery / safety attitudes of the S/C …

Introduction / EPS general information EPS equipments glossary Equipment

Equipment

Battery Charge Regulator

BCR

Maximum Peak Power Tracking

Battery Discharge Regulator

BDR

Power Conditioning Unit

Begin Of Life

BoL

Power Conditioning & Distribution Unit

Converter

CV

Power Subsystem

MPPT PCU PCDU PSS

Depth Of Discharge

DoD

Regulaterd bus

RB

End of Charge

EoC

Sequential Switching / Shunt Regulator

S3R

End of Discharge

EoD

Solar Array

End Of Life

EoL

Solar Array Drive Mechanism

Electrical Power Subsystem

EPS

State of Charge

Fold-back Current Limiter

FCL

Solid State Power Controller

Latchning Current Limiter

LCL

Unregulated Bus

39 Ref :

VLR/080066

Date :

21.10.13

SA SADM SoC SSPC URB

Introduction / EPS general information EPS in practice (1) TELECOM 2 (1991) Telecommunication – P = 3 kW

TELECOM 2

Mass (kg)

Satellite mass ratio

Satellite dry mass

1100

100 %

43

4%

Distribution

21

2%

Battery NiH2

132

12 %

Solar Array

100

9%

Power TOTAL

296

27 %

Power System

(incl. SADM)

Data from CNES

40 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information EPS in practice (2) JASON 1 (2001) Mini satellite Oceanographic Observation satellite – P = 500 W Mass (kg)


Satellite dry mass

472

100 %

Power System

10

2%

Distribution

29

6%

Battery NiCd

45

10 %

Solar Array

42

9%

Power TOTAL

126

27 %

Jason 1

(incl. SADM)

Data from CNES

Illustration CNES 41 Ref :

VLR/080066

Date :

21.10.13

Introduction / EPS general information EPS in practice (3) DEMETER (2004) Micro Satellite Science (Geodesy) satellite – P = 110 W Mass (kg)


Satellite dry mass

110

100 %

Power System

6.5

6%

4

4%

Solar Array

6.5

6%

Power TOTAL

17

16 %

DEMETER

(incl. SADM)

Battery LiIon

Data from CNES

Illustration CNES

42 Ref :

VLR/080066

Date :

21.10.13

Agenda 1. Introduction n Presentation of Thales Alenia Space Belgium n Presentation of myself n EPS – general information





43 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays SA cells n A solar cell is composed of a semiconductor material and converts photons to electrons n Photovoltaic effect n The solar flux is reflected, absorbed by the solar cell or crosses it n Every absorbed photon whose energy is greater than semiconductor 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 44 Ref :

VLR/080066

Date :

21.10.13

Illustration SPECTROLAB

Primary power sources / Solar cells & arrays Semi-conductors properties (1) n Most common semi-conductors: silicium and arsenide-gallium (AsGa) n Cubic crystalline structures

Silicium

n

Arsenide Gallium

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

• • •

45 Ref :

VLR/080066

Date :

21.10.13

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

Primary power sources / Solar cells & arrays Solar flux n Semi-conductor gap is chosen to fit with space light wavelength

46 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays 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 I

47 Ref :

VLR/080066

Date :

21.10.13

V

Primary power sources / Solar cells & arrays Optimal working point – at max. available power U-I characteristic 6.0

Isc

Pmax

5.0

I sa (A)

4.0 3.0

Psa characteristics 2.0

Psa-BoL-Hot

1.0

240 220

Isa-BoL-Hot

Pmax

Psa-EoL-Hot

Psa-BoL-Cold 200 18056 60 64 68 72 76 80 84 Isa-BoL-Cold 4 8 12 16 20 24 28 32 36 40 44 48 52 160 U sa (V) 140 Isa-EoL-Hot

0.0

P sa (W)

0

Pmax is largely depending of temperature & ageing

Voc

120 100 80 60 40 20 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) 48 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Efficiency

Illustration SPECTROLAB 49 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Efficiency degradation factors n Mission lifetime n Loss of power: 1% to 2% every year (depends of the orbit) n Radiation effects

n UV Illustration SPECTROLAB n Meteorite impact n 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)

50 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Concentrators / Advantages n Concentrate SA flux on SA cells n Reduce SA cells surface n Based on reflectors or on lens

Direct Illumination

51 Ref :

VLR/080066

Date :

21.10.13

Reflected Illumination

Primary power sources / Solar cells & arrays Concentrators / Drawbacks n 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/C n Induces higher thermal constraints on cells and SA panel Concentration at large off-pointing 2

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

1.8 1.6 1.4 Co nce 1.2 ntr ati 1 on 0.8

Temperature map 0.08

120

110

0.06

100 0.04

90 0.02

80 0 70

0.6

-0.02 60

0.4 -0.04

50

0.2 -0.06 -0.08

0 0

10

20

30

40

50

Off-Pointing (arcdeg)

n Outgassing may become critical 52 Ref :

VLR/080066

Date :

21.10.13

60

70

-0.06

-0.04

-0.02

0

0.02

0.04

0.06

Primary power sources / Solar cells & arrays 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 parallel n 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 independent Illustration Thales Alenia Space

53 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar arrays (2)

Illustration Thales Alenia Space

Protection against shadowing

54

Illustration SPECTROLAB Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar arrays (3) – Types n Fixed n Solar cells are glued on the structure of the satellite n The power is limited by the surface of the satellite n 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 mobile n 1-degree of freedom

55 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar arrays (3) – Types

56 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Panel, glue, coverglass, … n Substrate n Kapton with glass – or carbon- reinforcement n Glue, Adhesive n Fix SA cell on SA panel n Fix the coverglass on the cell n Ensure electrical & thermal conductivity n Panel (honeycomb) n Support SA cells n Transfer heat to bottom side n Face high thermal gradient n Be compatible with deployment and orientation mechanisms n Coverglass n Protect SA cell against ATOX n Protect SA cell against radiation n 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) 57 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar Arrays (4) – mechanisms (holddown & release) n SA are fold during launch n Deployment is n Initiated by pyro actuation (or thermal knifes) n Controlled by the use of hinge mechanisms

Illustration Thales Alenia Space 58 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar Arrays (4) – mechanisms (orientation) n Mobile SA is controlled by SADM n Current is transferred to S/C main part via BAPTA n Tensioning wires – achieve minimum fundamental frequency of the array (AOCS constraints)

Illustration Thales Alenia Space 59 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar arrays (5) – thermal interfaces n SA cell temperature directly impacts its efficiency n Temperature is linked to n The incoming flux

n

n

• Direct solar flux • Albedo • IR flux of the earth The outcoming flux • Flux reflected by the cells • Power delivered to the satellite • IR flux of the front and rear part of the SA 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)

60 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Solar cells & arrays Solar arrays (6) – performances n Typical performances after 15 years in GEO n 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/kg n Multi junctions: 50-60 W/kg

61 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Fuel cells Electromechanical devices performing a controlled chemical reaction (oxidation) to derive electrical energy (rather than heat energy) n Advantages n Minimal thermal changes n Compact and flexible solution n Production of water (manned mission) n Drawbacks n Need of fuels: hydrogen & oxygen yielding water as the reaction product n Used for Shuttle orbiter, lunar rover, …

62 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / Fuel cells Typical current-voltage curve for a hydrogen/oxygen fuel cell

Performance summary of fuel cells for space use System

Specific power (W/kg)

Operation

Comment

Gemini Apollo

33 25

240 h

Not drinking water

Shuttle

275

2500 h

110 – 146 367 110 550

> 40000 h > 3000 h > 40 000 h

SPE technology Alkaline technology Alkaline technology Goal (lightweight cell) 63 Ref :

VLR/080066

Date :

21.10.13

Operated at 505 K 24 h start-up / 17 h shutdown 15 min start-up / instantaneous shutdown

Primary power sources / Fuel cells 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 Advantage n Lower SA power need thanks to judicious sizing of the fuel n Drawback n 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

64 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / RTG Deep-space missions (further than Mars) or Military use n Long time missions, not-compatible with fuel cells n 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 effect n Generation of a voltage between (semi-conductor) materials maintaining a temperature difference. Power function of: n Absolute t° of hot junction n T° difference between materials n 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 65 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / RTG Advantages n n n n n

Power production independent of S/C orientation & shadowing Independence of distance from Sun Low power level may be provided for long time period Not susceptible to radiation damage Compatible with long eclipse (e.g. lunar landers)

Drawbacks n 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 66 Ref :

VLR/080066

Date :

21.10.13

Primary power sources / RTG & others Example of RTG n n n n

Cassini (Saturn mission) Galileo probe/Ulysses Nimbus/Viking/Pionner Appolo lander

n Mars Science Laboratory

628 W 285 W 35 W 25 W 73 W 120 W

195 W/kg 195 W/kg 457 W/kg 490 W/kg 261 W/kg 416 W/kg

Nuclear fission n 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

67 Ref :

VLR/080066

Date :

21.10.13

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 station n Reduced drag (reducing area of SA panels) n Reduced maintenance effort

68 Ref :

VLR/080066

Date :

21.10.13






69 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Accumulators n Electromechanical devices performing a controlled chemical reaction to derive electrical energy n Critical parameters n Charge/discharge rate n Depth of Discharge n Extent of over-discharging n Thermal sensitivity to each of these parameters

70 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Accumulators n Elementary Battery model 5mΩ

200 kF 3V3

Typical Characteristics: Cell voltage versus DoD & I discharge

Cell voltage (V)

n Capacity n 1.5Ah -> 100Ah n Voltage range n 4.1V -> 3.3V n Series Resistance n 1mΩ -> 10m Ω n Leakage current n 0mA -> 5mA

4,2 4,1 4,0 3,9 3,8 3,7 3,6 3,5 3,4 3,3 3,2 3,1 3,0 2,9 2,8

C C/2 C/5

0

5

10

15

20

25

30

Capacity (Ah)

71 Ref :

VLR/080066

Date :

21.10.13

35

40

45

50

Secondary power sources Comparison of performances Type

NiCd

NiH2

Li-Ion

Energy/kg (Wh/kg)

30-40

55-65

100-130

Energy/l (Wh/l)

110

80

200-250

Discharge voltage mean (V)

1.25

1.25

3.5

[-5;+15]

[0;+10]

[+15;+25]

⇒ C/10 (GEO) ⇒ C/2 (LEO)

⇒ C/8 (GEO) ⇒ 0.7C (LEO)

C/10 ⇒ C/3

⇒ 2C

⇒C

⇒C

75

75

90

Max. voltage (A)

1.55

1.6

4.0

Min. voltage (A)

1.0

1.0

2.7

4 ⇒ 50

30 ⇒ 350

1.5,2.2,26,40

Life duration in Geo

7 years at 50 % of DoD



Life duration in Leo




Working temperature (°C) Charge current (A) Discharge current (A) Energy efficiency

Capacity (Ah)

72 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Battery n Role: support the Solar Array during n LEOP phases n Eclipses n Loss of sun pointing n Peak power demands n … n Series / parallel assembling of accumulator cells n In series to reach the desired voltage

• • • •

n

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

In parallel to reach the desired capacity

Illustration SAFT

Illustration SAFT Illustration SAFT

73 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Battery failure modes n Open circuit n Loss of battery n Short failure n Degradation of the voltage Open circuit

Short failure

NiCd

Never met

Yes

NiH2

Yes

Yes

Li-Ion

Yes ? Not very well known

Yes ? Not very well known

74 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Battery protections n By-pass n 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-time n Counterbalancing of cells mismatching

75 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Battery accommodation n Mechanical & thermal drivers n The structure shall be rigid (launch environment) n The structure shall allow the homogeneity of the temperature

ZONE 2b - LOWER RADIAL PANEL 10000

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

Acceleration [g]

1000

100

10

1 1

10

76 Ref :

VLR/080066

Date :

21.10.13

100 Frequency [Hz]

1000

10000

Secondary power sources Battery sizing

77 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Sampling of accumulators SAFT NiCd VOS 40

SAFT NiH2 93 AN

Capacity

46 Ah

89 Ah

38.6 Ah

6.1 Ah

1.4 Ah

Mean voltage

1.2 V

1.36 V

3.6 V

3.6 V

3.7 V

Energy

55 Wh

120 Wh

140 Wh

22 Wh

5.2 Wh

Mass

1610 g

2108 g

1107 g

155 g

41.2 g

Energy/kg

34 Wh/kg

57 Wh/kg

126 Wh/kg

141 Wh/kg

126 Wh/kg

Efficiency

70 %

70 %

90 %

90 %

90 %

BoL

SAFT LiIon SAFT LiIon VOS140 MP76065

Data CNES

78 Ref :

VLR/080066

Date :

21.10.13

SONY LiIOn 18650HC

Secondary power sources Sampling of batteries SPOT5

EURASIASAT

STENTOR

SKYBRIDGE

24s-1p VOS40

27s-1p 93AN

11s-2p VOS140

12s-4p VOS140

Capacity

46 Ah

93 Ah

80 Ah

154 Ah

Mean voltage

29 V

37 V

39.6 V

43 V

1325 Wh

3415 Wh

3168 Wh

6620 Wh

47.4 kg

66 kg

34 kg

72 kg

467X261X260mm

863X441X310mm

490X380X290mm

910X520X300mm

28 Wh/kg

52 Wh/kg

93 Wh/kg

92 Wh/kg

42 Wh/l

29 Wh/l

59 Wh/l

47 Wh/l

BoL Configuration

Energy Mass Dimensions Specific energy Density

Data CNES

79 Ref :

VLR/080066

Date :

21.10.13

Secondary power sources Sampling of batteries BoL Configuration Capacity Mean voltage Energy Mass Dimensions Specific energy Density

PLEIADES

µSAT

8s-100p 18650HC

8s-10p 18650HC

140 Ah

14 Ah

30 V

30 V

4200 Wh

420 Wh

40,4 kg

4 kg

2 x (355 x 295 x 180 mm)

226 x 166 x 95 mm

105 Wh/kg

105 Wh/kg

170 Wh/l

118 Wh/l

Data CNES

Illustration SAFT

80 Ref :

VLR/080066

Date :

21.10.13






81 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / Architecture Block diagram Divided in sections

Bus SA

PCDU Distribution

User's

SA conditioning

DET, MPPT, …

BAT

BCR / BDR / Switches

Monitorings

1 or 2 batteries

-> Operation with primary & secondary power sources whose characteristics are changing with time and conditions of operations 82 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / Architecture Bus voltage selection Divided in sections

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

Bus SA

n Choice is based on n Bus power

• •

n

-> Some architecture may even requires two buses !!

Ref :

VLR/080066

Date :

21.10.13

Distribution SA conditioning

DET, MPPT, …

BAT

1 or 2 batteries


Monitorings

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, …)

Payload flight heritage

83

PCDU User's

Power management, control & distribution / Architecture Conditioning architecture n 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 phase

n Unregulated bus n Bus voltage is imposed by the battery voltage

•

Impact on all DC/DC converters efficiency

n Semi-regulated bus n Regulated bus in sunlight only n Choice is based on n User’s need (mission)

• •

n

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)

User’s flight heritage

84 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / Architecture Conditioning topology The 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 SA n 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 on n Mission & orbit

• •

n

…

85 Ref :

VLR/080066

Date :

21.10.13

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

Power management, control & distribution / Architecture Battery management n Architecture n Centralized

• •

n

Performed by OBC PCDU functions limited to monitoring

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 extremes n Choice based on Divided

• • •

Price Satellite reliability need …

in sections

SA SA conditioning

DET, MPPT, …

1 or 2 batteries Ref :

VLR/080066

Date :

21.10.13

PCDU Distribution

BAT

86

Bus


Monitorings

User's

Power management, control & distribution / Architecture Battery management n Functions n Voltage / current control during charge n Current limitation during discharge n Monitoring n Protections

87 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / Architecture Distribution architecture Distribution 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 Fuse n

Simplest solution n Imposes all the user’s to be compatible with bus transients induces by fuse blowing n Imposes the need of extraction during AIT phase n SSPC n Flexible solution PCDU Bus Divided n ON/OFF switching capability in SA User's Distribution sections n Control of fault current SA conditioning

DET, MPPT, …

BAT

1 or 2 batteries

88 Ref :

VLR/080066

Date :

21.10.13


Monitorings

Power management, control & distribution / Architecture Distribution architecture / some definitions Current falls depends on loads characteristics (L / C) only

I

n LCL n n n n

Imax < 1.2 x Ilim.max

Latching Current Limiter Limits current at user’s switch ON or short failure during limitation time Trips-OFF if limitation time is exceeded ON/OFF command capability

Ilim.max Ilim.nom Ilim.min

Inom dI/dt < 1 A/µs

n FCL Fold-back current limiter n 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 LCL n Essential load (e.g. OBC) n LCL + automatic periodic re-arming n … n

t Trip-off time

Reaction time < 1µs

FCL I-V CHARACTERISTICS Voltage

Iclass Inom

Vbus

Ilimit-min

Ilimit-max

Current

89 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / Architecture Other constituents of PCDU n Command of mechanisms n SADM motor driver n Antenna motor driver n … n Command on deployment n Actuation of pyro n Actuation of thermal knifes n Li-Ion battery cells management n Acquisition of thermistors n …

90 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / Architecture Battery management / Li-Ion battery cells management n Cells balancing n Voltage balancing at cells level via the actuation of a shunt in parallel with the cell to slightly discharge it n 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

91 Ref :

VLR/080066

Date :

21.10.13

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 converter n Not reliable n Distribution functions n LCL, Pyro n DC/DC for secondary (+5, +-15,+20 V) + LCL protection n Adaptability of the distribution by paralleling of LCL’s n CNES/Astrium/TAS-F Myriade platform baseline

92 Ref :

VLR/080066

Date :

21.10.13

Illustration CNES

Power management, control & distribution / PC(D)U Examples Scientific & earth observation n n n n n n n n n n

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

93 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / PC(D)U Examples Scientific & earth observation

Illustration CNES

Illustration ESA 94 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / PC(D)U Examples Geo low power SPACEBUS 3000 PCU n Full regulated bus 5.5 kW / 50 V n Solar array regulation: S3R n No distribution function (PCU only) n Flight heritage : 29 PCU’s, 200 years

95 Ref :

VLR/080066

Date :

21.10.13

Power management, control & distribution / PC(D)U Examples Geo high power SPACEBUS 4000 PCU n Full regulated bus 6 to 21 kW / 100 V n Solar array regulation: S3R n No distribution function (PCU only) n Flight heritage : 13 PCU’s, 300 000 h

96 Ref :

VLR/080066

Date :

21.10.13






97 Ref :

VLR/080066

Date :

21.10.13

Power budget Study case n Study of a micro satellite to target ship based and ground based radars n Lifetime: 5 years n Orbit: Leo n Payload requirements n Acquisition in sun & eclipse phases n Bus power of 626 W

• • • • •

98 Ref :

VLR/080066

Date :

21.10.13

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

Power budget Orbit selection n Altitude trade-off n Lower than 1000 km (to avoid Van Allen belts impacts on radiation level n 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 worldwide n Sun-synchronous orbit as other candidate Orbit characteristics

99 Ref :

VLR/080066

Date :

21.10.13

Average height Period Eccentricity Inclination

600 km 5820 s 0.001 (circular orbit) 90 ° (polar orbit)

Eclipse duration

21.3 min

600 km 5820 s 0.001 (circular orbit) 98 ° (sunsynchronous) 30 min

Power budget Orbit selection / Inclination trade-off 1500 1400

1200


1000


1000

800

600

500 400

200

0

0

200

400

600

800

1000

1200

1400

1600

days

1800

0

0

200

400

600

800

1000

1200

1400

days

SA flux (sun-synchronous orbit)

SA flux (polar orbit)



100 Ref :

VLR/080066

Date :

21.10.13

1600

1800

Power budget Orbit selection / Inclination trade-off n 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).

Minimum SA flux (W/m2) BOL SA cell efficiency EOL/BOL ratio Total available SA power (W / m2)

Sun-synchronous

Polar

1220

520 28 % 76.5 %

261

Cell manufacturer data

111

n Note that photovoltaic efficiency EOL/BOL ratio takes into account the following elements (SA panel manufacturer data) n 5-years mission lifetime n radiation effects n UV and meteoritic impact n effect of ATOX density (aggressive and corrosive environment tied to the LEO) on cover glass protection

101 Ref :

VLR/080066

Date :

21.10.13

Power budget EPS sizing / bus voltage trade-off n 28 V n

Compatible with bus power (< 1 kW) n High hardware heritage n 50 V n Reduced current levels n Reduced harness & power dissipations

EPS siring / Battery sizing & bus regulation trade-off n Regulated power bus – main hypothesis n BDR (Battery => bus) conversion efficiency n Unregulated power bus – main hypothesis n Internal losses (Battery => bus) internal connections n Battery n Max DOD of 40 % considered following

• •

Orbit characteristics (period and eclipse) Mission duration

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

102 Ref :

VLR/080066

Date :

21.10.13

92% 1%

Power budget EPS sizing / bus regulation trade-off (sun-synchronous orbit) & Battery sizing Regulated bus Users' power in eclipse (W) PCDU losses during eclipse (W) Satellite power requirement in eclipse (W) Eclipse duration (min)

84 710

Pout/n – Pout + LL

Unregulated bus 626 36 662 30

(Pecl*30 min)/Tcharge Battery useful cycled energy requirement (W) Maximum acceptable DoD Battery energy requirement (Wh) @ 40% DoD Battery energy requirement (Wh) @ 40% DoD with 10% margin

355

331 40%

888

828

977

911

Energy/DOD

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

PCDU BAPTA

S3R S4R

regulated 28V

PDU

BAPTA

MPPT cell

unregulated 28V

PDU

BDR

power distribution to users

Li-Ion Battery

103 Ref :

VLR/080066

Date :

21.10.13

power distribution to users

Li-Ion Battery

Power budget EPS sizing / Conditioning topology trade-off n Unregulated topology n n

Internal PCU losses (Battery => bus) PCDU low level consumption

1% 30 W for both configurations

n MPPT n n

Conversion efficiency 95 % Ability to track the maximum power whatever the battery state is (charged, discharged, with or without failure, …).

n DET n n n

Conversion efficiency 97.5 % S4R (SA => battery) conversion efficiency 95.5% 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)

n Battery data (based on previous selection) n

n

Psa characteristics

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

• •

25 W (discharge) 15 W (charge)

PCU to battery harness losses = 3 %

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

Psa-BoL-Hot 240 220

P sa (W)

n

Psa-EoL-Hot Psa-BoL-Cold

200 180 160 140 120 100 80 60 40 20 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)

104 Ref :

VLR/080066

Date :

21.10.13

Power budget EPS sizing / Conditioning topology trade-off MPPT

DET 662 30 25 1.005 15 357 11

Satellite power requirement in eclipse (W) Eclipse duration (min) Discharge dissipation losses (W) Battery recharge duration (h) Recharge dissipation losses (W) Battery recharge power requirement (W) 3% Precharge Battery - PCU harness losses (W) S4R / MPPT losses (W) SA power requirement for battery recharge (W)

18 386

Users' power in sunlight (W) PCDU losses during sunlight (W) Satellite power requirement in sunlight (W)

63 689

46 672

TOTAL SA requirepment (W) TOTAL SA requirepment (W) with 10% margin

1075 1182

1056 1162

SA efficiency (W/m2) Minimum SA surface requirement (m2)

261 4.5

235 4.9

(97 -30)*0.9 ((Pecl+Lossbat)*30 min)/Tcharge + Lossbat

17 384 626

5 or 4.5% Pbus charge including harness losses

Pout/n – Pout + LL

Non-optimization : 10%

MPPT offers about 9 % SA surface reduction (compared with DET solution) & greater flexibility towards satellite FDIR but imposes some constraints on thermal interfaces (dissipation inside PCDU is higher) & impacts PCDU definition (more complex solution) Sun-synchronous Polar 105 Ref :

VLR/080066

Date :

21.10.13

Minimum SA flux (W/m2) BOL SA cell efficiency EOL/BOL ratio Total available SA power (W / m2)

1220

520 28 % 76.5 %

261

111






106 Ref :

VLR/080066

Date :

21.10.13

Conclusions

The design of any Power Subsystem is strongly linked with System analyses (Attitude & Orbit, Mission, Operations) The electrical architecture of spacecrafts is not standard n n n n n n

Unregulated or regulated Power bus Voltage (28 V, 50 V, 100 V, …) Conditioning (S3R, MPPT, …) Protections (reliable or not) Distribution (fuse, LCL, …) …

and shall be adapted nearly on case by case ….

107 Ref :

VLR/080066

Date :

21.10.13

Thales Alenia Space ETCA n n n n

Tél. : + 32/71 44 22 11 Fax : + 32/71 44 22 00 www.thalesaleniaspace.com E-mail : [email protected]

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

108

10/04/2007 Ref :

VLR/080066

Date :

21.10.13