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Lithium Ion Battery Management Strategies for European Space Operations Centre Missions Thomas Ormston * , Viet Duc Tran , Michel Denis and Nic Mardle § European Space Agency, Darmstadt, Germany Luke Lucas and Laurent Maleville LSE Space GmbH, Darmstadt, Germany Kees Van Der Pols Telespazio VEGA Deutschland GmbH, Darmstadt, Germany Effective battery management on a space mission is one of the key factors in ensuring mission success and longevity. Given the reliability of modern spacecraft, the unavoidable ageing of batteries can become a critical life-limiting factor. To improve this, it is necessary to have a strategy for management and monitoring of spacecraft batteries that is tailored to both the mission profile and the battery technology in use. This paper will focus on several missions flown from the European Space Agency’s (ESA) European Space Opera- tions Centre (ESOC) in Darmstadt, Germany. The main case studies in this paper focus on missions that regularly use their Lithium Ion batteries, although a summary of other missions that contain Lithium Ion batteries will also be presented. Lithium Ion batteries are currently the prevailing battery technology in use on current and future European Space Agency missions. The paper will begin with an overview of the Lithium Ion battery technology that has largely replaced all others for modern space batteries. Their proper management requires different techniques compared to previous space battery technologies; for instance compared to the previous Nickel-Cadmium technology, Lithium Ion battery deep discharges should be avoided where possible - which increases the risk of using deep discharges to measure degradation. The paper will describe the characteristics and influencing factors of Lithium Ion battery degradation, along with an overview of research aimed at prolonging lifetime of the batteries. The paper will also summarise methods available in order to measure the absolute or relative degradation of Lithium Ion batteries and the limitations of these methods based upon the capabilities of each spacecraft and the mission profile. The paper will then detail the actual operational implementation of this information on two representative ESA missions. The first case study will be Mars Express, which has been flying three Lithium Ion batteries for ten years and using them for prolonged eclipse seasons 2-3 times per year. The power demand of the spacecraft is high and the available margin in the power system is low, therefore modelling and management of the batteries is critical to the mission. The second case study will be ESA’s CryoSat-2 mission, which has been flying a single Lithium Ion battery for 4 years. The battery is younger, and the * Spacecraft Operations Engineer, Earth Observation Missions - EarthCARE (HSO-OER), European Space Operations Cen- tre, 64293 Darmstadt Germany. Spacecraft Operations Engineer, Earth Observation Missions - Aeolus (HSO-OEA), European Space Operations Centre, 64293 Darmstadt Germany. Spacecraft Operations Manager, Planetary Missions - Mars Express (HSO-OPM), European Space Operations Centre, 64293 Darmstadt Germany. § Spacecraft Operations Manager, Earth Observation Missions - CryoSat-2 (HSO-OEE), European Space Operations Centre, 64293 Darmstadt Germany. Spacecraft Operations Engineer, Planetary Missions - Mars Express (HSO-OPM), European Space Operations Centre, 64293 Darmstadt Germany. Spacecraft Operations Engineer, Earth Observation Missions - SWARM (HSO-OEW), European Space Operations Centre, 64293 Darmstadt Germany. 1 of 18 American Institute of Aeronautics and Astronautics Downloaded by 189.137.147.147 on July 7, 2015 | http://arc.aiaa.org | DOI: 10.2514/6.2014-1883 SpaceOps 2014 Conference 5-9 May 2014, Pasadena, CA AIAA 2014-1883 Copyright © 2014 by European Space Agency. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. SpaceOps Conferences

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  • Lithium Ion Battery Management Strategies for

    European Space Operations Centre Missions

    Thomas Ormston, Viet Duc Tran, Michel Denis and Nic Mardle

    European Space Agency, Darmstadt, Germany

    Luke Lucas and Laurent Maleville

    LSE Space GmbH, Darmstadt, Germany

    Kees Van Der Pols

    Telespazio VEGA Deutschland GmbH, Darmstadt, Germany

    Eective battery management on a space mission is one of the key factors in ensuringmission success and longevity. Given the reliability of modern spacecraft, the unavoidableageing of batteries can become a critical life-limiting factor. To improve this, it is necessaryto have a strategy for management and monitoring of spacecraft batteries that is tailoredto both the mission prole and the battery technology in use. This paper will focus onseveral missions own from the European Space Agency's (ESA) European Space Opera-tions Centre (ESOC) in Darmstadt, Germany. The main case studies in this paper focuson missions that regularly use their Lithium Ion batteries, although a summary of othermissions that contain Lithium Ion batteries will also be presented. Lithium Ion batteriesare currently the prevailing battery technology in use on current and future EuropeanSpace Agency missions.

    The paper will begin with an overview of the Lithium Ion battery technology thathas largely replaced all others for modern space batteries. Their proper managementrequires dierent techniques compared to previous space battery technologies; for instancecompared to the previous Nickel-Cadmium technology, Lithium Ion battery deep dischargesshould be avoided where possible - which increases the risk of using deep discharges tomeasure degradation. The paper will describe the characteristics and inuencing factors ofLithium Ion battery degradation, along with an overview of research aimed at prolonginglifetime of the batteries. The paper will also summarise methods available in order tomeasure the absolute or relative degradation of Lithium Ion batteries and the limitationsof these methods based upon the capabilities of each spacecraft and the mission prole.

    The paper will then detail the actual operational implementation of this information ontwo representative ESA missions. The rst case study will be Mars Express, which hasbeen ying three Lithium Ion batteries for ten years and using them for prolonged eclipseseasons 2-3 times per year. The power demand of the spacecraft is high and the availablemargin in the power system is low, therefore modelling and management of the batteriesis critical to the mission. The second case study will be ESA's CryoSat-2 mission, whichhas been ying a single Lithium Ion battery for 4 years. The battery is younger, and the

    Spacecraft Operations Engineer, Earth Observation Missions - EarthCARE (HSO-OER), European Space Operations Cen-tre, 64293 Darmstadt Germany.

    Spacecraft Operations Engineer, Earth Observation Missions - Aeolus (HSO-OEA), European Space Operations Centre,64293 Darmstadt Germany.

    Spacecraft Operations Manager, Planetary Missions - Mars Express (HSO-OPM), European Space Operations Centre,64293 Darmstadt Germany.

    Spacecraft Operations Manager, Earth Observation Missions - CryoSat-2 (HSO-OEE), European Space Operations Centre,64293 Darmstadt Germany.

    Spacecraft Operations Engineer, Planetary Missions - Mars Express (HSO-OPM), European Space Operations Centre,64293 Darmstadt Germany.

    Spacecraft Operations Engineer, Earth Observation Missions - SWARM (HSO-OEW), European Space Operations Centre,64293 Darmstadt Germany.

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    SpaceOps 2014 Conference 5-9 May 2014, Pasadena, CA

    AIAA 2014-1883

    Copyright 2014 by European Space Agency. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

    SpaceOps Conferences

  • power system has more margin but eclipse seasons are an almost constant feature of theroutine mission (albeit with varying duration eclipses). In addition, the satellite ies in anon-sun-synchronous orbit, which makes the assessment of the expected state of batterycharge more dicult. An overview of the techniques used on other ying ESOC missionswill also be presented (Herschel, Planck, GOCE, Venus Express and Rosetta).

    The paper will describe new operations that have been introduced to manage the degra-dation of the batteries, including specially designed settings that, while respecting the al-lowed usage prole of the battery, modify the charge and discharge management strategiesand other ight operations to almost halve the rate of degradation compared with theworst-case design assumption. In addition, the methods used by each mission to assessabsolute and/or relative battery degradation in ight will be discussed.

    The paper will conclude with an overview of the lessons that have been learnt so far atESOC from missions ying Lithium Ion batteries. These lessons could be used as a modelfor current and future operators of spacecraft with Lithium Ion batteries on how to bestmanage their batteries for longevity, mission reliability and success.

    I. Introduction

    The majority of ESA missions now use Lithium Ion batteries as their primary method of power storage.This battery technology has many advantages for spacecraft and mission design and is foreseen for manyof the upcoming missions to be own by ESOC too. However, although they are an accepted and welcomeadvancement for spacecraft design, the operational use of Lithium Ion batteries in ight is still a relativelynew area.

    It is becoming increasingly common that spacecraft will survive long past their design life and as suchthe proper operational management of Lithium Ion batteries from day one of the mission, including prior tolaunch, is critical. This eectively requires two components - determining the level of battery degradationand reducing the rate of battery degradation.

    At present the strategies for this management have been developed largely as a benet of ight experience,and often independently from one another due to the dierent generations of missions and the dierentdemands and constraints of various missions. This paper aims to summarise Lithium Ion battery technologyas a primer to operators working with the technology and to further go on to highlight some diering butrepresentative examples of Lithium Ion battery management at ESOC. It will conclude with a summary ofsome lessons learned during the time ESOC has been operating spacecraft with Lithium Ion batteries.

    II. Space Battery Technology

    For the majority of spacecraft, whether scientic, communication or other purpose, in LEO, MEO, GEOor interplanetary, power and its supply at all times is a matter of careful consideration.

    During sunlight illumination power is normally provided by solar arrays, however spacecraft in all ofthe orbits mentioned will experience solar eclipses and need a secondary power source for these periods.Rechargeable batteries have been used for this purpose as they are particularly applicable to long durationmissions. Non-rechargeable energy sources e.g. primary batteries for short missions or radioisotope ther-moelectric generator power sources for deep space missions are not within the scope of this paper. Whilerechargeable batteries have been a common part of spacecraft systems since early in the space age, thebattery chemistry has undergone signicant changes.

    A battery is made up of a number of cells in series and parallel. The arrangement of the cells providesthe required current and voltage. The voltage of any cell or battery depends upon the electrochemistry ofthe cell. It is here that changes in the cell electrochemical make-up have brought about great changes inspace batteries. This electrochemistry impacts the key parameters of a battery, namely:

    Capacity - Dened as the number of Ampere Hours (Ah) a battery can deliver at room temperatureuntil it reaches a cut-o voltage where it can no longer deliver power routinely (commonly two thirdsof the fully charged voltage); it is dependent upon the size of the battery.

    Specic Energy - The energy stored per unit mass and measured in Watt Hours per Kilogram(Wh/kg). Reduced mass of the battery for the same power storage allows a heavier and/or morecomplex payload.

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  • Energy Density - The energy stored per unit volume and measured in Watt Hours per Litre (Wh/l).Reduced volume of the battery for the same power storage has a similar positive impact on spacecraftdesign.

    Cycle Life - The number of charge/discharge cycles a battery can undergo and still provide theminimum required voltage.

    Although not a parameter of the battery itself, but rather a parameter of the state of a battery, theDepth of Discharge (DoD) is dened as the Ah capacity drained from battery, divided by the real Ahcapacity of the battery at that point (i.e. including any measured or assumed degradation of the battery).This measure is also sometimes referred to as its inverse, the State of Charge (SoC), equal to 1 minusthe DoD. The SoC refers to the level of charge remaining in the battery. Both of these measures are oftenreferred to as percentages, achieved by multiplying the value by 100.

    A. Space Battery Technology Evolution

    The rst common space batteries in the 1960s were Nickel Cadmium (Ni-Cd). They became the coretechnology of many missions and were the most common battery in use up to the mid-1980s. They are wellcharacterised and known. The space industry rides on the back of the tried and tested Ni-Cd battery and itis still powering spacecraft like XMM-Newton today.

    A Ni-Cd battery is temperature sensitive and has a typical specic energy of approximately 25 Wh/kg.Ni-Cd has a high cycle life which is important for long duration missions; however, signicantly the Ni-Cdbattery is subject to the `Memory Eect'. The Ni-Cd battery `remembers' its most frequently used DoDand does not work well beyond that. That is to say, a Ni-Cd battery frequently discharged to 25% DoD willnot be able to discharge lower than 25% DoD, therefore eectively losing capacity and rendering the batteryunable to provide the required power to the spacecraft. The potential loss of capacity can be circumventedby regularly `re-conditioning' the battery by discharging it nearly completely. While this operation restoresthe battery it requires additional hardware on-board the spacecraft and additional periodic operations ofthe spacecraft in orbit. More recently, cadmium has also evoked environmental concerns and is subject toregulatory scrutiny.

    In the years since the rst space batteries, such as the early ESA mission ESRO-2 in 1968, improvementshave been made in electrochemistry and newer cell types have been developed. In the 1980s, Nickel Hydrogen(Ni-H2) batteries, a successor to Ni-Cd, came into use. Ni-H2 cells combine and make the best of two dierentchemistries - that of the Nickel Oxide electrode of the Ni-Cd cell and that of the Hydrogen catalyst electrodeof the fuel cell.

    Their chemistry allows a deeper DoD for a comparable cycle than Ni-Cd, resulting in a lower requiredAh capacity which in turn leads to lower mass. As the cost of space ight depends upon mass, any savingis highly advantageous.

    Ni-H2 cells have been widely used in both LEO and GEO spacecraft. The evolved chemistry lead to ahigher reliability and longer lifetime in orbit compared to Ni-Cd, while avoiding the penalty of regulatedcadmium.

    In 2001 a move away from nickel-based batteries and to a new chemistry began with the ight of therst Lithium Ion (Li-ion) battery aboard Proba-1, an ESA technology demonstrator. The Li-ion cell oers asignicant leap in performance. They oer a high specic energy of 85-130 Wh/kg, a 3-5 fold improvementin specic energy compared to Ni-Cd cells. A comparison of mass and volume of Ni-Cd, Ni-H2 and Li-ionbatteries can be seen in gure 1. The huge gains in mass and volume oer new design possibilities forincreasing payload mass, volume and/or power demand. In addition the modular concept of Li-ion batteriesgives benets of simplicity while also allowing exibility in accommodation which was exploited in ESAmissions such as Mars Express, CryoSat-2 and Philae.

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  • ... ..Ni-Cd

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    (b) Battery Volume Comparison

    Figure 1. Mass and volume comparison of Lithium Ion (Li-ion, carbon cells) against heritage battery tech-nologies. Values given are for a 10 kWh battery with a maximum DoD of 75%.1

    The Li-ion cell is also magnetically `cleaner' than the nickel batteries, which can be signicant in sensitiveinstrumentation. Aside from the spacecraft itself, the handling of Li-ion is also considerably easier. Priorto launch the Li-ion battery is very easy to store and has a long storage life. As many programs experiencedelays and launch slips, this pre-launch factor is non-negligible. After launch, once in operation the Li-ionbattery does not suer from any `memory eect'. This makes the battery easier to operate and does awaywith the need for additional hardware for the `re-conditioning' process and the added complexity of theoperation itself.

    B. Space Battery Technology Usage

    Li-ion batteries are now ying on or have own on LEO, GEO and Interplanetary missions. As their spaceheritage becomes established, their performance has proven to be even better in ight than predicted. Mis-sions are able to y longer and experience more cycles at higher DoD while continuing to achieve substantialpayload results.

    Examples of space batteries in ESA spacecraft and the evolution of battery chemistry is shown in table1 below.

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  • Table 1. Summary of battery type and mission type for a selection of ESA missions.

    MissionName

    MissionPurpose

    LaunchDate

    FlightDuration

    Orbit BatteryType

    NumberofBatteries

    BatterySize

    ESRO-2 Science 17.05.68 3 years LEO Ni-Cd 1 3 Ah

    COS-B Science 09.08.75 7 years HEO Ni-Cd 1 6 Ah

    Marecs-A Telecom 20.12.81 14.75 years GEO Ni-Cd 2 21 Ah

    Giotto Science 02.07.85 7 years DeepSpace

    Ag-Cd 4 16 Ah

    Olympus Telecom 12.07.89 4 years GEO* Ni-CdNi-H2

    11

    24 Ah35 Ah

    ERS-1 EarthObservation

    17.07.91 5 years LEO Ni-Cd 4 24 Ah

    Eureca Microgravity 31.07.92 1 year LEOy Ni-Cd 4 40 AhXMM-Newton Science 10.12.99 Ongoing HEO Ni-Cd 2 24 Ah

    Proba-1 TechnologyDemonstration

    22.10.01 Ongoing LEO Li-ion 1 9Ah

    Envisat EarthObservation

    01.03.02 10 years LEO Ni-Cd 8 40 Ah

    MSG-1 Weather 28.08.02 Ongoing GEO Ni-Cd 2 29 Ah

    Integral Science 17.10.02 Ongoing HEO Ni-Cd 2 24 Ah

    Mars Express Science 02.06.03 Ongoing Planetary Li-ion 3 22.5 Ah

    SMART-1 Science 28.09.03 3 years Lunar Li-ion 5

    Rosetta Science 02.03.04 Ongoing DeepSpace

    Li-ion 3 16.5 Ah

    MSG-2 Weather 21.12.05 Ongoing GEO Ni-Cd 2 29 Ah

    Venus Express Science 09.11.05 Ongoing Planetary Li-ion 3 24 Ah

    GOCE EarthObservation

    17.03.09 4.5 years LEO Li-ion 1 78 Ah

    Herschel Science 14.05.09 4 years L2 Li-ion 1 39 Ah

    Planck Science 14.05.09 4.5 years L2 Li-ion 1 39 Ah

    CryoSat-2 EarthObservation

    08.04.10 Ongoing LEO Li-ion 1 78 Ah

    *Batteries were used for LEOP and for 2 major recoveries. Both batteries were frozen due to a spacecraft anomaly, thensuccessfully recovered to operation.

    yRetrieved by Space Shuttle and returned to Earth at end of mission.

    This evolution of ESA missions to the use of Li-ion batteries is thanks to the benets of the technology.Figure 2 summarises the battery types of the table above and graphically shows how space battery technologyhas evolved over time as greater energy density technologies have become available.

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  • 1950 1970 1990 2010

    Energy Density

    Silver-Zinc

    Nickel-Cadmium/ Argon-Zinc

    Nickel- Hydrogen

    Lithium-Ion

    Sputnik 1 1956

    PROBA 2001

    Mars Express 2003

    CryoSat-2 2010

    Sentinel Family 2014

    Bepi Colombo 2016

    Figure 2. Evolution of the energy density and use in the space eld of battery technologies over time.

    C. Lithium Ion Lifetime Management

    Despite the demonstrated advantages of Li-ion batteries for space use, they do still degrade over time and assuch this must be managed to ensure that the optimal possible lifetime of the batteries is achieved. Lifetimemanagement can only be addressed after dening the health descriptors of the batteries. Battery healthindicators at the highest level are:

    Remaining Capacity - The same as the Capacity mentioned earlier, but adjusted to give the realamount of Ah that the battery can deliver. This is governed by two types of degradation; degradationby discharge cycles and degradation due to ageing.

    Internal Resistance - The internal resistance of the cells in the battery. This is key to the performanceof the battery - higher currents in or out cause a loss of resulting energy that can be delivered. Theincrease of internal resistance is predominantly governed by temperature, but also by ageing.

    Battery Health

    Internal ResistanceRemaining Capacity

    Capacity fadedue to ageing

    Capacity fadedue to cycling

    Figure 3. The constituent parts that make up the health of a battery.

    Lifetime management for Lithium Ion batteries aims to minimise the possible impacts from the types ofdegradation discussed above. The key to minimising degradation is to manage and mitigate where possiblethe degradation drivers given below.

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  • Capacity Fade Due To Cycling

    The capacity fade due to cycling is primarily impacted by the DoD reached on a given cycle and thetotal number of cycles experienced by the battery. Deeper and/or more frequent discharges will leadto a higher capacity fade rate.

    Capacity Fade Due To Ageing

    The capacity fade due to ageing cannot be avoided, however storing the batteries colder and not fullycharged signicantly helps keeping the ageing rate low. The capacity fade due to ageing is quicker athigher temperatures. Li-ion batteries should always be kept below 20 degC and if possible close to 0degC. 0 to 5 degC seems to be the optimum range for Li-ion battery longevity. A reduction of thisageing of the batteries can be realised by abstaining from continuously trickle charging the batteriesonce they have reached 100% SoC.

    Stress factors due to (dis)charging at high rates are also good to avoid, as this keeps internal resistanceduring cycling inside a low range.

    All degradation types described above are more pronounced in the rst phases of the lifecycle due to theexponential character of most eects.

    III. Case Study 1: Mars Express

    A. Power System Design Description

    The Power subsystem on Mars Express consists of two solar array wings and three batteries, both connectedto a Power Conditioning Unit by means of Array Power Regulators (APRs) and Battery Charge/DischargeRegulators (BCDRs), respectively. The Power Conditioning Unit (PCU) takes the input of these powersources and regulates the power to a 28V bus. The Power Distribution Unit (PDU) then distributes this28V regulated bus to the various users on the spacecraft. The PCU has three operating modes, dependingon the availability of the power sources versus the power demand of the spacecraft:

    APR Mode

    Spacecraft Power Demand + Battery Charge Demand Maximum Possible Solar Array Supply

    In this case the maximum solar array supply still exceeds the amount required by the spacecraft, butis not enough to reach the maximum battery charge rate. In this mode the solar array is operatedat maximum power output, supplying the whole requirements of the spacecraft power bus, and anyremaining power being routed to charge the battery, meaning that charging rate varies depending onthe instantaneous level of unused solar array power.

    BDR Mode

    Spacecraft Power Demand (alone) >Maximum Possible Solar Array Supply

    In this case the solar arrays are operated at maximum power output but this is still not enough tosupply the spacecraft demand. In this case the batteries are discharged to make up the shortfall inavailable power. A typical example of this case would be during a spacecraft eclipse.

    In the specic case of Mars Express there is a design anomaly which means that the harness connectingthe solar cells to the APRs is incomplete, causing a lower power performance than was designed. Thewiring harness between the Solar Arrays and the APRs is such that the Maximum Power Point Trackingvoting system cannot function as was foreseen. At best 72% of the design power is available from the solararrays. This reduced power was and is a restrictive issue for Mars Express. Nonetheless, the mission has

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  • outperformed the dened scientic goals in terms of lifetime extension and scientic return due to carefulmanagement of the power system and accurate planning tools.

    In terms of the batteries used on Mars Express, they consist of three 24 Ah (at Beginning of Life) Li-ionbatteries. Each battery is composed of multiple Sony Cells manufactured into space qualied hardware byABSL Space Products.

    B. Battery Usage Prole

    On Mars Express, the battery usage is dened by two factors - the eclipse seasons and augmenting the poweravailable from the arrays during special operations or low-power seasons. The key driver for the degradationof the battery are the large and frequent discharges caused by eclipse seasons. These seasons can last anumber of months with typically a short gap in between. In the rst years of the mission, the eclipses werelonger than in recent years, with maximum eclipse duration reaching 90 minutes in 2004, reducing to a peakeclipse duration during a season of 40 minutes in 2014. This aligns well with the state of the power system,as the demands placed on the batteries are reduced as they age. The degradation rates of the battery havenot been as high as anticipated, and as such there is power margin to work with and extended discharges (i.e.on top of eclipse demand for science pointings as well as outside eclipse seasons) are allowed and relativelycommon. This trade-o allows extended science objectives to be realised by allowing certain pointings thatimpact the DoD following an eclipse. An example of such a \double discharge" case is given in this graph,going to 30% DoD twice in one orbit.

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    Figure 4. Mars Express battery DoD over approximately one orbit on DoY 2014-074. First peak is due tosolar eclipse by Mars, the second due to science pointing that required suboptimal array pointing.

    The eclipse seasons on Mars Express cover approximately 60% of the mission life, and during this timethere will be at least one discharge/charge cycle every orbital revolution. Throughout the years, being ableto more accurately plan and monitor discharges and their degradation impact has been critical to ensure

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  • proper management of the batteries as a limited resource available to the mission.

    C. Lifetime Preservation Measures

    As the Mars Express batteries are indeed a limited resource, it is important to take any necessary measuresto reduce the rate of their degradation. Following the information from the previous chapter, 5 key principlesfor reduction in degradation rate were identied:

    Minimise depth of discharge

    In the case of eclipses, there is no way to avoid the discharge, but we try to make sure the DoD isonly as high as strictly required. In the orbit, the most interesting part for most science observationsis around closest point to Mars (pericentre). The eclipses take place just before pericentre or overlapwith it, from which two power-related problems arise. During eclipse, if payload is switched on, theDoD will be higher as a result of more power demand. After eclipse, a special pointing might berequired to enable the science observation and the sun aspect angle to the arrays cannot be optimised,leading to extra discharge outside of eclipse. Minimising DoD comes down to a trade-o between thevalue of performing a science observation which eectively enlarges the DoD and on the other handsafeguarding the longevity of the Mars Express mission by preserving the batteries. Mars Expresshas mission planning rules (max routine DoD = 45%) and resource allocation processes in place toaddress requests for higher DoDs in a structured manner. Special power optimised pointings have beendeveloped to assist in this process. The transmitter is o by default in eclipse and not switched onuntil well after the eclipse to create a solid recharge margin. Also, in some mission seasons, the largestheater groups are phased to use most power outside of eclipse by pre-eclipse boost heating.

    Minimise number of cycles

    Minimising the number of cycles is achieved by making sure that, outside of eclipses, no excess power isdemanded from the solar arrays requiring a battery discharge, unless strictly planned for and allowedduring the mission planning process. This is also the reason for not performing too many BatteryCapacity Measurements in ight, since they have a signicant impact based on high DoD and extradischarge/charge cycle.

    Store at correct temperature

    The temperature at which Mars Express batteries are nominally operated is between -5 and 0 degC.While the batteries do have heaters to ensure they do not get too cold, they are largely above thistemperature and the driver for their temperature is the steady state temperature of the Mars Expressspacecraft platform, rather than specic battery thermal control. During discharges, temperatureexcursions of +10-15 degC are seen, depending on the DoD and discharge rates. This can be seenbelow in gure 5.

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    Figure 5. Mars Express battery temperature from 2010 until 2014. Raised periods correspond with batteryin use during eclipse seasons.

    Store at lower state of charge

    It is known that a lower charge level will result in a slower degradation rate due to ageing. On top ofthat, it has been proven in ight on Mars Express that reducing the nominal state of charge (wheneverpossible outside of eclipse seasons) from 100% down to 80-90% had a positive eect on the availablecapacity by the time the new eclipse season was due to start. In gure 6 we can see that during eclipseseasons, the discharge started from 100% SoC, and outside eclipse seasons the SoC was lowered (therewere still some discharges in this period but these were unavoidable and analysed to ensure they weresmall and safe).

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    Figure 6. Mars Express battery SoC from 2010 until 2014. Inter-eclipse seasons show extended periods wherebattery end of charge level was lowered to 80-90% for lifetime preservation purposes.

    Keep charge and discharge rates low

    The discharge rate is controllable based on which platform and payload equipment is on and by designthe charge rate is set to the maximum that can be delivered by the arrays after subtracting thespacecraft bus requirements. This is up to a charge rate limit of 9 Amps (but the BCRs can also beset to 3 Amps). Since the solar array harness to the APRs is incomplete, the output from the arrays isless, but from the point of view that slower charge is better for battery preservation, this has a positiveeect here. The regular charge rates lie between 1 and 4 Amps depending on the mission season.

    D. Capacity Measurement

    The analysis of the power budget of Mars Express only covered the originally planned mission duration of 1Martian year plus a second Martian year as an extension (totalling approximately 4 Earth years). Estimateson battery capacity degradation therefore only cover this period and take a very conservative predictionof the degradation rate. To be able to assess the health state of the batteries more accurately, and togather trends for predictive models, the Flight Control Team had to develop its own empirical methodsbased on data available from telemetry. A test under laboratory conditions is not possible due to severalfactors: no calibration of the sensing equipment is possible in-ight, no possibility to remove the batteryfrom the circuit, no possibility to control the discharge rate and no possibility to ne-control the thermalenvironment. The method used here was initiated by the Mars Express ight control team with assistancefrom the ESA battery technology experts and further developed and ne-tuned by the Venus Express ightcontrol team. The model used consists of tting a battery voltage curve (based on the Beginning of Life`State-of-Charge vs. EMF' curve) to the voltage measurements as obtained from telemetry. The modeloptimises for initial energy capacity degradation factor and the losses due to internal resistance dissipation.

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  • The internal resistance modelling is variable to account for (dis)charge rates as well as thermal eects (heat-up during discharge and cool-down during charge). Of particular interest is the successful modelling of thetransition between discharge and charge. The model is covered in more detail in Ref. 2.

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    Figure 7. Battery voltage for Venus Express Deep Discharge Test 9, performed on DoY 2013-285. The batteryvoltage can be seen to drop past the critical level where a faster rate of discharge begins. The model tting,used to assess degradation, can be seen to t well with the telemetry data.

    To obtain a complete dataset for this analysis the operators had to plan for dedicated deep discharges,as data available from routine eclipses was not sucient to obtain a signicant excursion of the EMF curveoutside its linear initial part. The discharge down to 60-65% DoD showed the portion of the curve wherethe steeper second part started, the exact position of which is indicative of the battery degradation level.Example telemetry and model tting of such a test is shown in gure 7. These dedicated Deep DischargeTests were performed in visibility (for safety reasons) and were triggered by rotating the Solar Arrays edge-on to the Sun. The visibility requirement causes the discharge rate to be relatively high as the X-bandtransmitter is on. On-board protection was added to ensure that at a given battery voltage the arrays wouldrotate back to the normal position even in case of loss of ground contact.

    Both Mars Express and Venus Express ight control teams have determined long term trends of thecapacity evolution (which is useful for long term mission planning and approval of mission extensions) andused it for short term prediction of the battery voltage behaviour for activities where the power marginavailable is critical. Establishing a better relation between SoC and voltage allowed the ight control teamto be able to use the margin, when needed, more safely. Any allowed DoD gure in the mission planningcycle can be coupled to the corresponding voltage used and remaining capacity.

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  • IV. Case Study 2: CryoSat-2

    A. Power System Design Description

    The power system of CryoSat-2 is driven by the constraints of its non-sun-synchronous orbit, which featuresa wide variability in sun incidence angle on the spacecraft and periods of eclipse where no solar array poweris available.

    The CryoSat-2 Power Subsystem (EPS) provides the following functions:

    Generation of electrical power by means of a solar array

    Control, storage and distribution of electrical power to/via a main bus

    Battery management (charge/discharge/protection)

    Provision of unregulated main bus power to the units attached to this bus in the range of 22 to 34 V

    Provision of status monitoring and telecommand interfaces for subsystem operation and performanceevaluation

    Provision of adequate redundancy and protection circuitry to avoid failure propagation and to ensurerecovery from any malfunction within the subsystem and/or load failure.

    The CryoSat-2 EPS comprises the following units:

    A two panel xed (non-rotating) GaAs Solar Array with 11 electrical sections per panel, regulated bymeans of a shunt system

    A single 78 Ah battery, consisting of 52 strings with 8 Li-ion cells each, manufactured by ABSL SpaceProducts.

    Combined Power Control and Distribution Unit

    Battery charge is controlled by the PCDU in order to prevent the battery from experiencing thermalstress. During insucient solar array power, either due to eclipse or high demand, for example in a dawn-dusk orbit with sub-optimal array alignment, the energy stored in the battery will be used to satisfy thebus power demand. The Battery is charged by applying an IV-method with charge current as the controlparameter besides battery voltage. Eight commandable end of charge (EoC) voltage limits (steps of 200mV) allow compensation of any potential cell parameter variations of the battery.

    B. Battery Usage Prole

    The CryoSat-2 orbit is a polar orbit but it is not sun-synchronous and consequently the orbital plane rotateswith respect to the sun direction. The nodal plane regresses at a rate of about 0.25 per day. It thereforemakes half a revolution, sampling all local solar times, in just over 8 months. This means that the satellitefaces great variations in solar illumination and there are periods when it ies along the dawn-dusk line andis in constant sunlight, but with only one of the solar arrays illuminated. At other periods it ies in thenoon-midnight plane with both solar arrays illuminated and undergoes eclipses. At maximum extent theeclipse duration is around 36 minutes.

    This leads to seasons of battery use, as with Mars Express, at the present state of the spacecraft. However,as the arrays age, the poor illumination in the dawn-dusk seasons could lead to more regular battery use tocompensate for array output.

    C. Lifetime Preservation Measures

    Although the CryoSat-2 mission is younger and has more margin in its power system, it is still consideredimportant to take measures to reduce the rate of battery degradation in order to preserve the resource ofbattery life for future use. As for Mars Express, CryoSat-2 obeys the 5 key principles of Li-ion battery lifepreservation:

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  • Minimise depth of discharge

    This is performed \de facto" on CryoSat-2, rather than actively managed as with Mars Express. This isthanks to the greater design margins in the CryoSat-2 power system and the more stable and repetitivenature of its power demand. This allows the operators to ensure that no excessive discharges will occuras long as routine operations are conducted.

    Minimise number of cycles

    As with Mars Express, the cycles from eclipse periods are unavoidable. However, outside of eclipseperiods, extraneous cycles are largely avoided by the design margins, as in the previous point.

    Store at correct temperature

    The batteries on CryoSat-2 are maintained by thermal control to a region above 10 degrees. This is stillmostly low enough to prevent excessive capacity fade degradation. The increased temperature (withrespect to Mars Express) prevents the increase in internal resistance at lower temperatures, which hastwo drawbacks:

    { Lower charge/discharge eciency, as the internal losses are higher

    { Lower charge capability, as the end of charge voltage limit is reached already at a lower SoC, dueto the higher voltage drop

    Store at lower state of charge

    During the early days of a LEO type mission, batteries are often being charged to 100% SoC evenwhen the missions DoD is around 30%. This unnecessary margin in the power system has been usedon CryoSat-2 to reduce the degradation due to ageing by minimising the SoC of the battery during thewhole mission. For example, there could be no energy demand from the battery for several weeks everyfew months when the spacecraft is in a dawn-dusk orbit. During these battery non-operational periodsthe SoC of the battery is reduced in order to reduce the ageing degradation. However, it is important toensure that the reduction in SoC during mission does not interfere with the energy demand during fulloperation. During the transition between no eclipse to longest eclipse the battery voltage at the endof the discharge is monitored and when it crosses a predened threshold the end of charge is modiedsuch as to always have enough power available to ensure the safety of the satellite even in a worst caseanomaly, while maintaining a low (around 80%) SoC at the end of the charge whenever possible.

    Keep charge and discharge rates low

    This is another area where CryoSat-2 has passive control, with low charge/discharge rates being ensuredby the spacecrafts conservative design margins rather than by active operator control.

    D. Capacity Measurement

    CryoSat-2 faces similar issues to Mars Express in that measuring internal battery characteristics such as theinternal resistance and the battery open circuit voltage is not achievable in ight. In fact the CryoSat-2housekeeping telemetry provides us only with the battery charge/discharge current and the battery terminalvoltage, along with a summary of these in the on-board software. In addition, the solar arrays have a xedposition and so cannot be turned away to provoke a deep discharge of the battery, as with Mars Express.Therefore estimations of the battery degradation parameters have to be performed with routine ight data.

    1. Battery internal resistance estimation

    One of the best ways in ight to estimate the battery internal resistance is to induce a high impulsivecurrent demand and measure the corresponding terminal voltage drop, a method detailed in the CryoSat-2User Manual.3 This method works well if the current demand is large but the current demand on CryoSat-2is typically quite low, with the largest peaks being around 3 amps, for example when a heater is switchedon. However, even with this low current the method can still be applied to see if trends on the evolution ofthe battery internal resistance appear over satellite lifetime.

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  • 2. Battery Capacity Estimation

    To estimate capacity of the battery without deep discharges, the CryoSat-2 User Manual3 proposed a methodbased on comparing the voltage drop and integrated current produced by the battery. Using the theoreticalcurve for a new battery which converts voltage into state of charge provided by ABSL, one can get thestate of charge change between these two points and then estimate the capacity by computing what a 100%discharge will produce. This method is similar to the deep discharge method used on Mars Express but doesnot provide reliable results for CryoSat-2 due to the relatively small discharges.

    Instead it has been preferred to monitor the degradation of the battery by comparing the expected Ahconverted from the voltage drop seen in TM using the ABSL curve and the eective Ah measured by telemetryfor each longest eclipse. The ABSL curve provides a conversion between voltage level of the battery andstate of charge and can be used therefore to produce an estimate of the Ah that a given voltage drop shouldproduce for a battery with a given capacity. Periodically we compare the power that should be produced by abattery with a capacity as estimated at the previous longest eclipse with the power produced now computedby integrating the current for the latest longest eclipse. This method is similar to the previous method but,while not capable of providing an absolute value of remaining capacity, it does provide a relatively accuraterelative value of capacity lost between tests.

    Using the percentage of reduction between the power produced at the previous longest eclipse and theone at the last eclipse measurement gives us a factor to apply to the last estimated capacity of the batteryto estimate the current capacity of the battery. This iteration was started using the capacity given by ABSLfor the battery at beginning of life.

    The results of this method are shown in gure 8 below and show a reasonable match with the predictionmade by ABSL.

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    Figure 8. Evaluated vs. predicted battery capacity degradation for the CryoSat-2 battery. The manufacturer(ABSL) predicted degradation of capacity matches well with the battery capacity degradation that has beenevaluated using telemetry.

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  • V. Summary of Li-ion Battery Management on other ESOC Missions

    As mentioned previously, ESA's Proba-1 mission, launched in 2001, was the rst spacecraft to orbit Earthwith commercial Li-ion battery technology on board. Since then almost all ESA missions have been equippedwith Li-ion batteries, gradually bringing to an end the era dominated by conventional nickel-cadmium andnickel-hydrogen batteries for spacecraft.

    The management and usage of Li-ion batteries diers from mission to mission depending on their orbitprole and power system design. As has already been mentioned, what all missions using Li-ion batterieshave in common are the important operational constraints for maximum charge and discharge voltagesand currents, correct settings for end of charge levels and the avoidance of thermal stress. All this iscommonly taken into account and managed by the PCDU in combination with respective thermal regulationwhich by design optimises the battery lifetime and reduces the operational overhead. Furthermore the levelof operational experience has increased with time and missions, providing a set of applied best-practicescontributing to the battery longevity, and extending missions considerably as highlighted in the above casestudies. In the process of justifying the technical and operational feasibility to extend an ESOC mission,Li-ion batteries have therefore proven to be an essential factor.

    To complement the above case studies, the following table gives an overview of other ESOC missions andtheir respective battery management characteristics. While the above case studies apply to missions withpower systems designed to cope with regular eclipse encounters or, in case of Mars Express, operations waybeyond their nominal lifetime, other missions exist where lithium-ion batteries are not used in routine ightat all or have shown to require no additional operational measures. For example, Herschel and Planck, eachequipped with a 36 Ah Li-ion battery, were ying in an orbit without eclipses where together with the solararray design and spacecraft orientation sunlight was permanently available to supply the bus load. Hence,the battery was only used during launch when no solar array power was available. Given that Lithium-ionbatteries do not require reconditioning and are free of memory eects, no activities for capacity preservationor capacity measurements were needed throughout the mission.

    GOCE had a 78 Ah battery at the front of the spacecraft which operated awlessly throughout the mis-sion, encountering seasons of 16 eclipses per day with peak durations up to 32 minutes. Except for staying inthe dened boundaries in which the battery had to be operated no particular battery management strategieshad to be applied due to the mission life-limiting factor being propellant, not battery life. Battery perfor-mance assessment using in-ight telemetry to compute the battery capacity spent during discharge cyclesshowed good performance of the batteries. A battery simulation on ageing and number of charge/dischargecycles based on in-ight telemetry conducted by industry in preparation for the low orbit operations in 2012concluded a battery degradation due to ageing and cycling of about 6%, whereas 19.5% was expected atthat time. Shortly before the re-entry the battery was successfully operated at temperatures above 80 degCfor a short time.

    The Rosetta mission is a similar case to Herschel/Planck where the batteries are not used in routine.Although not foreseen, the three 16.5 Ah batteries were used for a Mars y-by in February 2007 when thespacecraft entered solar-eclipse and for solar array performance tests in 2010 and 2014 (before and after deepspace hibernation) where a step-wise sun o-pointing of the solar arrays made use of the batteries to supplythe required power and assess how much power the solar arrays could provide. To preserve the batterycapacity the end-of-charge level was lowered to 87%.

    Venus Express on the other hand is similar to Mars Express where the extension of the mission required adetailed understanding of the degradation level and solar eclipse seasons could only be survived with the on-board battery. To determine the battery capacity Deep Discharge tests through solar array o-pointing wereperformed, as with Mars Express. The collected data was tted to dedicated models from which predictionson the capacity degradation and internal resistance trend could be made. Furthermore, the end-of-chargelevel was lowered outside eclipse seasons to 80%. Having kept the batteries in a good shape will be crucialin the 2014 period where the batteries will be used to allow extended periods of low aspect angles on thearrays and high DoDs during aerobraking in the Venusian atmosphere.

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  • Table 2. Summary of Li-ion battery monitoring and management strategy for ESOC missions.

    Mission Battery Usage Prole Lifetime PreservationMeasures

    Capacity MeasurementIn-Flight

    GOCE Launch, eclipses and peakpower demands

    None Performed In-ight telemetry usedfor battery simulationtests

    Venus Express Launch, eclipses and peakpower demands

    Battery end-of-chargelevel lowered to 80% SoC

    Deep discharge tests andMonte-Carlo tting

    Rosetta Launch, Mars swing-by inFeb. 2007 (eclipse) andSolar Array performancetests in 2010 and 2014

    Battery end-of-chargelevel lowered to 80% SoC

    None performed

    Herschel/Planck Launch only None Performed None Performed

    VI. Lessons Learned

    Throughout this process lessons have been learned that could contribute to future or current operatorsworking with Lithium Ion batteries.

    1. All operational lifetime preservation measures presented in this paper are useful, even though indi-vidually they may seem minor. Each measure, when applicable, allows to slow down the capacitydegradation by up to a few percent over mission lifetime with respect to the spacecraft manufacturer'sprediction.

    2. The mission, spacecraft and power system design heavily inuence the ability to perform lifetimepreservation measures. Operators should work with the spacecraft designers to ensure as far as possiblethat the design allows the exibility to perform these measures.

    3. Operators should also work with the spacecraft designers to ensure that sucient telemetry is availablein ight to accurately monitor battery health. This could include increase of sensitivity of sensors,number of sensors or frequency of samples.

    4. Ground and space system modelling of the SoC and DoD, either in prediction, planning or telemetry,should take into account a variable degradation factor and allow this factor to be changed easily inorder to match the real degradation.

    5. The most eective (and only absolute) way to reliably measure true degradation of a Lithium Ion spacebattery in ight is through a Deep Discharge Test.

    6. The ability to perform a Deep Discharge Test should be included in the design of spacecraft powersystems, even if they normally would not have that ability (i.e. CryoSat-2). The mechanism used todo this would have to be carefully considered to ensure it could be performed at negligible risk to themission.

    7. Measures to preserve and monitor Lithium Ion batteries should be implemented by operators as part ofthe operations concept of a mission, planned for before launch, to maximise their eectiveness. Whileit is never too late to implement such measures, they are most benecial if started from day one.

    8. A close working relationship between operations teams, spacecraft and power system designers andbattery experts is key to minimising the degradation and maximising the usage and lifetime of a spacebattery.

    VII. Conclusion

    This paper has reviewed some background on Lithium Ion space batteries and focussed on their use andmanagement on missions own from ESOC. Even in this small selection it is clear that the strategies being

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  • employed are variable. There are a number of reasons for this, as highlighted in the paper. These includemission design factors, such as the usage prole of the battery, the length of the mission and the hardwareavailable. However, the operational factors of how the battery is monitored and operated are also variedacross missions.

    Ultimately though, it is clear that there are a number of common threads that can be followed to bestmonitor Lithium Ion batteries and to reduce the rate of degradation. The lessons learned point to theneed to implement such strategies from the moment spacecraft operations start and to consider them in thespacecraft design and operations concept. It is also clear that there is still harmonisation to be done betweendierent operators of Lithium Ion batteries, spacecraft designers and battery experts. This will result in anoptimal realisation of the potential of this battery technology both in the spacecraft design phase and formany years of operations to come - maximising the potential and longevity of our space missions.

    Acknowledgments

    T. Ormston thanks all of the co-authors of this paper for their hard work and for producing an excellentoverview of Lithium Ion battery usage from across the spectrum of ESOC missions.

    The authors thank the battery technology experts from within ESA and from our industrial partnersthat have assisted us in building up our operational experience and knowledge of working with Lithium Ionbatteries.

    References

    1Dudley, G. and Verniolle, J., \Secondary Lithium Batteries for Spacecraft," ESA Bulletin, , No. 90, May 1997, pp. 50{54.2Sousa, B. and Van Der Pols, C. L., \Breath in, breath out, how healthy are the batteries on Mars and Venus Express,"

    SpaceOps, Stockholm, Sweden, June 2012.3Astrium GmbH, ., \CryoSat-2 User Manual," CS-MA-DOR-SY-0001.

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    IntroductionSpace Battery TechnologySpace Battery Technology EvolutionSpace Battery Technology UsageLithium Ion Lifetime Management

    Case Study 1: Mars ExpressPower System Design DescriptionBattery Usage ProfileLifetime Preservation MeasuresCapacity Measurement

    Case Study 2: CryoSat-2Power System Design DescriptionBattery Usage ProfileLifetime Preservation MeasuresCapacity MeasurementBattery internal resistance estimationBattery Capacity Estimation

    Summary of Li-ion Battery Management on other ESOC MissionsLessons LearnedConclusion