Lithium Ion Battery Management Strategies for European Space Operations Centre Missions

Transcription

1 SpaceOps Conferences 5-9 May 2014, Pasadena, CA SpaceOps 2014 Conference / 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 Operations 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 Centre, Darmstadt Germany. Spacecraft Operations Engineer, Earth Observation Missions - Aeolus (HSO-OEA), European Space Operations Centre, Darmstadt Germany. Spacecraft Operations Manager, Planetary Missions - Mars Express (HSO-OPM), European Space Operations Centre, Darmstadt Germany. Spacecraft Operations Manager, Earth Observation Missions - CryoSat-2 (HSO-OEE), European Space Operations Centre, Darmstadt Germany. Spacecraft Operations Engineer, Planetary Missions - Mars Express (HSO-OPM), European Space Operations Centre, Darmstadt Germany. Spacecraft Operations Engineer, Earth Observation Missions - SWARM (HSO-OEW), European Space Operations Centre, Darmstadt Germany. 1 of 18 Copyright 2014 by European Space Agency. Published by the American American Institute of Institute Aeronautics ofand Aeronautics Astronautics, Inc., andwith Astronautics permission.

2 power system has more margin but eclipse seasons are an almost constant feature of the routine mission (albeit with varying duration eclipses). In addition, the satellite flies in a non-sun-synchronous orbit, which makes the assessment of the expected state of battery charge more difficult. An overview of the techniques used on other flying ESOC missions will also be presented (Herschel, Planck, GOCE, Venus Express and Rosetta). The paper will describe new operations that have been introduced to manage the degradation of the batteries, including specially designed settings that, while respecting the allowed usage profile of the battery, modify the charge and discharge management strategies and other flight operations to almost halve the rate of degradation compared with the worst-case design assumption. In addition, the methods used by each mission to assess absolute and/or relative battery degradation in flight will be discussed. The paper will conclude with an overview of the lessons that have been learnt so far at ESOC from missions flying Lithium Ion batteries. These lessons could be used as a model for current and future operators of spacecraft with Lithium Ion batteries on how to best manage 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 many of the upcoming missions to be flown by ESOC too. However, although they are an accepted and welcome advancement for spacecraft design, the operational use of Lithium Ion batteries in flight is still a relatively new area. It is becoming increasingly common that spacecraft will survive long past their design life and as such the proper operational management of Lithium Ion batteries from day one of the mission, including prior to launch, is critical. This effectively requires two components - determining the level of battery degradation and reducing the rate of battery degradation. At present the strategies for this management have been developed largely as a benefit of flight experience, and often independently from one another due to the different generations of missions and the different demands and constraints of various missions. This paper aims to summarise Lithium Ion battery technology as a primer to operators working with the technology and to further go on to highlight some differing but representative examples of Lithium Ion battery management at ESOC. It will conclude with a summary of some lessons learned during the time ESOC has been operating spacecraft with Lithium Ion batteries. II. Space Battery Technology For the majority of spacecraft, whether scientific, communication or other purpose, in LEO, MEO, GEO or 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 of the 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 duration missions. Non-rechargeable energy sources e.g. primary batteries for short missions or radioisotope thermoelectric generator power sources for deep space missions are not within the scope of this paper. While rechargeable batteries have been a common part of spacecraft systems since early in the space age, the battery chemistry has undergone significant changes. A battery is made up of a number of cells in series and parallel. The arrangement of the cells provides the required current and voltage. The voltage of any cell or battery depends upon the electrochemistry of the cell. It is here that changes in the cell electrochemical make-up have brought about great changes in space batteries. This electrochemistry impacts the key parameters of a battery, namely: Capacity - Defined as the number of Ampere Hours (Ah) a battery can deliver at room temperature until it reaches a cut-off voltage where it can no longer deliver power routinely (commonly two thirds of the fully charged voltage); it is dependent upon the size of the battery. Specific 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 more complex payload. 2 of 18

3 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 spacecraft design. Cycle Life - The number of charge/discharge cycles a battery can undergo and still provide the minimum required voltage. Although not a parameter of the battery itself, but rather a parameter of the state of a battery, the Depth of Discharge (DoD) is defined as the Ah capacity drained from battery, divided by the real Ah capacity 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 minus the DoD. The SoC refers to the level of charge remaining in the battery. Both of these measures are often referred to as percentages, achieved by multiplying the value by 100. A. Space Battery Technology Evolution The first common space batteries in the 1960s were Nickel Cadmium (Ni-Cd). They became the core technology of many missions and were the most common battery in use up to the mid-1980s. They are well characterised and known. The space industry rides on the back of the tried and tested Ni-Cd battery and it is still powering spacecraft like XMM-Newton today. A Ni-Cd battery is temperature sensitive and has a typical specific energy of approximately 25 Wh/kg. Ni-Cd has a high cycle life which is important for long duration missions; however, significantly the Ni-Cd battery is subject to the Memory Effect. The Ni-Cd battery remembers its most frequently used DoD and does not work well beyond that. That is to say, a Ni-Cd battery frequently discharged to 25% DoD will not be able to discharge lower than 25% DoD, therefore effectively losing capacity and rendering the battery unable to provide the required power to the spacecraft. The potential loss of capacity can be circumvented by regularly re-conditioning the battery by discharging it nearly completely. While this operation restores the battery it requires additional hardware on-board the spacecraft and additional periodic operations of the spacecraft in orbit. More recently, cadmium has also evoked environmental concerns and is subject to regulatory scrutiny. In the years since the first space batteries, such as the early ESA mission ESRO-2 in 1968, improvements have been made in electrochemistry and newer cell types have been developed. In the 1980s, Nickel Hydrogen (Ni-H 2 ) batteries, a successor to Ni-Cd, came into use. Ni-H 2 cells combine and make the best of two different chemistries - that of the Nickel Oxide electrode of the Ni-Cd cell and that of the Hydrogen catalyst electrode of the fuel cell. Their chemistry allows a deeper DoD for a comparable cycle than Ni-Cd, resulting in a lower required Ah capacity which in turn leads to lower mass. As the cost of space flight depends upon mass, any saving is highly advantageous. Ni-H 2 cells have been widely used in both LEO and GEO spacecraft. The evolved chemistry lead to a higher reliability and longer lifetime in orbit compared to Ni-Cd, while avoiding the penalty of regulated cadmium. In 2001 a move away from nickel-based batteries and to a new chemistry began with the flight of the first Lithium Ion (Li-ion) battery aboard Proba-1, an ESA technology demonstrator. The Li-ion cell offers a significant leap in performance. They offer a high specific energy of Wh/kg, a 3-5 fold improvement in specific energy compared to Ni-Cd cells. A comparison of mass and volume of Ni-Cd, Ni-H 2 and Li-ion batteries can be seen in figure 1. The huge gains in mass and volume offer new design possibilities for increasing payload mass, volume and/or power demand. In addition the modular concept of Li-ion batteries gives benefits of simplicity while also allowing flexibility in accommodation which was exploited in ESA missions such as Mars Express, CryoSat-2 and Philae. 3 of 18

4 Mass (kg) Volume (litres) Ni-Cd Ni-H 2.. Cells.. Battery Packaging (a) Battery Mass Comparison Li-ion Ni-Cd Ni-H 2.. Cells.. Battery Packaging (b) Battery Volume Comparison Li-ion Figure 1. Mass and volume comparison of Lithium Ion (Li-ion, carbon cells) against heritage battery technologies. 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 significant in sensitive instrumentation. Aside from the spacecraft itself, the handling of Li-ion is also considerably easier. Prior to launch the Li-ion battery is very easy to store and has a long storage life. As many programs experience delays and launch slips, this pre-launch factor is non-negligible. After launch, once in operation the Li-ion battery does not suffer from any memory effect. This makes the battery easier to operate and does away with the need for additional hardware for the re-conditioning process and the added complexity of the operation itself. B. Space Battery Technology Usage Li-ion batteries are now flying on or have flown on LEO, GEO and Interplanetary missions. As their space heritage becomes established, their performance has proven to be even better in flight than predicted. Missions are able to fly longer and experience more cycles at higher DoD while continuing to achieve substantial payload results. Examples of space batteries in ESA spacecraft and the evolution of battery chemistry is shown in table 1 below. 4 of 18

5 Table 1. Summary of battery type and mission type for a selection of ESA missions. Mission Name Mission Purpose Launch Date Flight Duration Orbit Battery Type Number Battery of Size Batteries ESRO-2 Science years LEO Ni-Cd 1 3 Ah COS-B Science years HEO Ni-Cd 1 6 Ah Marecs-A Telecom years GEO Ni-Cd 2 21 Ah Giotto Science years Deep Space Ag-Cd 4 16 Ah Olympus Telecom years GEO * Ni-Cd 1 Ni-H 2 1 ERS-1 Earth Observation 24 Ah 35 Ah years LEO Ni-Cd 4 24 Ah Eureca Microgravity year LEO Ni-Cd 4 40 Ah XMM-Newton Science Ongoing HEO Ni-Cd 2 24 Ah Proba-1 Technology Ongoing LEO Li-ion 1 9Ah Demonstration Envisat Earth years LEO Ni-Cd 8 40 Ah Observation MSG-1 Weather Ongoing GEO Ni-Cd 2 29 Ah Integral Science Ongoing HEO Ni-Cd 2 24 Ah Mars Express Science Ongoing Planetary Li-ion Ah SMART-1 Science years Lunar Li-ion 5 Rosetta Science Ongoing Deep Li-ion Ah Space MSG-2 Weather Ongoing GEO Ni-Cd 2 29 Ah Venus Express Science Ongoing Planetary Li-ion 3 24 Ah GOCE Earth years LEO Li-ion 1 78 Ah Observation Herschel Science years L2 Li-ion 1 39 Ah Planck Science years L2 Li-ion 1 39 Ah CryoSat-2 Earth Observation 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, then successfully recovered to operation. Retrieved 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 benefits of the technology. Figure 2 summarises the battery types of the table above and graphically shows how space battery technology has evolved over time as greater energy density technologies have become available. 5 of 18

6 Energy Density Silver-Zinc Nickel-Cadmium/ Argon-Zinc Nickel- Hydrogen PROBA 2001 Lithium-Ion Sentinel Family 2014 Sputnik Mars Express 2003 CryoSat Bepi Colombo Figure 2. Evolution of the energy density and use in the space field 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 as such this must be managed to ensure that the optimal possible lifetime of the batteries is achieved. Lifetime management can only be addressed after defining the health descriptors of the batteries. Battery health indicators at the highest level are: Remaining Capacity - The same as the Capacity mentioned earlier, but adjusted to give the real amount of Ah that the battery can deliver. This is governed by two types of degradation; degradation by discharge cycles and degradation due to ageing. Internal Resistance - The internal resistance of the cells in the battery. This is key to the performance of the battery - higher currents in or out cause a loss of resulting energy that can be delivered. The increase of internal resistance is predominantly governed by temperature, but also by ageing. Remaining Capacity Battery Health Internal Resistance Capacity fade due to cycling Capacity fade due to ageing 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 of degradation discussed above. The key to minimising degradation is to manage and mitigate where possible the degradation drivers given below. 6 of 18

7 Capacity Fade Due To Cycling The capacity fade due to cycling is primarily impacted by the DoD reached on a given cycle and the total number of cycles experienced by the battery. Deeper and/or more frequent discharges will lead to 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 fully charged significantly helps keeping the ageing rate low. The capacity fade due to ageing is quicker at higher temperatures. Li-ion batteries should always be kept below 20 degc and if possible close to 0 degc. 0 to 5 degc seems to be the optimum range for Li-ion battery longevity. A reduction of this ageing of the batteries can be realised by abstaining from continuously trickle charging the batteries once they have reached 100% SoC. Stress factors due to (dis)charging at high rates are also good to avoid, as this keeps internal resistance during cycling inside a low range. All degradation types described above are more pronounced in the first phases of the lifecycle due to the exponential character of most effects. III. A. Power System Design Description Case Study 1: Mars Express The Power subsystem on Mars Express consists of two solar array wings and three batteries, both connected to a Power Conditioning Unit by means of Array Power Regulators (APRs) and Battery Charge/Discharge Regulators (BCDRs), respectively. The Power Conditioning Unit (PCU) takes the input of these power sources and regulates the power to a 28V bus. The Power Distribution Unit (PDU) then distributes this 28V regulated bus to the various users on the spacecraft. The PCU has three operating modes, depending on 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 voltage at which power is taken from the solar arrays is backed off from the maximum power point so the actual power delivered by the arrays is equal to the spacecraft demand plus the battery charge demand. BCR 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, but is not enough to reach the maximum battery charge rate. In this mode the solar array is operated at maximum power output, supplying the whole requirements of the spacecraft power bus, and any remaining power being routed to charge the battery, meaning that charging rate varies depending on the 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 to supply the spacecraft demand. In this case the batteries are discharged to make up the shortfall in available power. A typical example of this case would be during a spacecraft eclipse. In the specific case of Mars Express there is a design anomaly which means that the harness connecting the solar cells to the APRs is incomplete, causing a lower power performance than was designed. The wiring harness between the Solar Arrays and the APRs is such that the Maximum Power Point Tracking voting system cannot function as was foreseen. At best 72% of the design power is available from the solar arrays. This reduced power was and is a restrictive issue for Mars Express. Nonetheless, the mission has 7 of 18

8 outperformed the defined scientific goals in terms of lifetime extension and scientific return due to careful management 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-ion batteries. Each battery is composed of multiple Sony Cells manufactured into space qualified hardware by ABSL Space Products. B. Battery Usage Profile On Mars Express, the battery usage is defined by two factors - the eclipse seasons and augmenting the power available from the arrays during special operations or low-power seasons. The key driver for the degradation of the battery are the large and frequent discharges caused by eclipse seasons. These seasons can last a number of months with typically a short gap in between. In the first years of the mission, the eclipses were longer than in recent years, with maximum eclipse duration reaching 90 minutes in 2004, reducing to a peak eclipse duration during a season of 40 minutes in 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 have not 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 relatively common. This trade-off allows extended science objectives to be realised by allowing certain pointings that impact 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. Depth of Discharge (%) :00 08:00 09:00 10:00 Time Figure 4. Mars Express battery DoD over approximately one orbit on DoY First peak is due to solar 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 time there will be at least one discharge/charge cycle every orbital revolution. Throughout the years, being able to more accurately plan and monitor discharges and their degradation impact has been critical to ensure 11:00 12:00 13:00 14:00 8 of 18

9 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 measures to reduce the rate of their degradation. Following the information from the previous chapter, 5 key principles for reduction in degradation rate were identified: 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 is only as high as strictly required. In the orbit, the most interesting part for most science observations is around closest point to Mars (pericentre). The eclipses take place just before pericentre or overlap with it, from which two power-related problems arise. During eclipse, if payload is switched on, the DoD will be higher as a result of more power demand. After eclipse, a special pointing might be required 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-off between the value of performing a science observation which effectively enlarges the DoD and on the other hand safeguarding the longevity of the Mars Express mission by preserving the batteries. Mars Express has mission planning rules (max routine DoD = 45%) and resource allocation processes in place to address requests for higher DoDs in a structured manner. Special power optimised pointings have been developed to assist in this process. The transmitter is off by default in eclipse and not switched on until well after the eclipse to create a solid recharge margin. Also, in some mission seasons, the largest heater 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 is demanded from the solar arrays requiring a battery discharge, unless strictly planned for and allowed during the mission planning process. This is also the reason for not performing too many Battery Capacity Measurements in flight, since they have a significant impact based on high DoD and extra discharge/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 this temperature and the driver for their temperature is the steady state temperature of the Mars Express spacecraft platform, rather than specific battery thermal control. During discharges, temperature excursions of degc are seen, depending on the DoD and discharge rates. This can be seen below in figure 5. 9 of 18

10 12 10 Battery 1 Battery 2 Battery Temperature (degc) Date Figure 5. Mars Express battery temperature from 2010 until Raised periods correspond with battery in 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 of that, it has been proven in flight on Mars Express that reducing the nominal state of charge (whenever possible outside of eclipse seasons) from 100% down to 80-90% had a positive effect on the available capacity by the time the new eclipse season was due to start. In figure 6 we can see that during eclipse seasons, the discharge started from 100% SoC, and outside eclipse seasons the SoC was lowered (there were still some discharges in this period but these were unavoidable and analysed to ensure they were small and safe). 10 of 18

11 State of Charge (%) Date Figure 6. Mars Express battery SoC from 2010 until Inter-eclipse seasons show extended periods where battery 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 design the charge rate is set to the maximum that can be delivered by the arrays after subtracting the spacecraft bus requirements. This is up to a charge rate limit of 9 Amps (but the BCRs can also be set to 3 Amps). Since the solar array harness to the APRs is incomplete, the output from the arrays is less, but from the point of view that slower charge is better for battery preservation, this has a positive effect 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 1 Martian year plus a second Martian year as an extension (totalling approximately 4 Earth years). Estimates on battery capacity degradation therefore only cover this period and take a very conservative prediction of the degradation rate. To be able to assess the health state of the batteries more accurately, and to gather trends for predictive models, the Flight Control Team had to develop its own empirical methods based on data available from telemetry. A test under laboratory conditions is not possible due to several factors: no calibration of the sensing equipment is possible in-flight, no possibility to remove the battery from the circuit, no possibility to control the discharge rate and no possibility to fine-control the thermal environment. The method used here was initiated by the Mars Express flight control team with assistance from the ESA battery technology experts and further developed and fine-tuned by the Venus Express flight control team. The model used consists of fitting 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 model optimises for initial energy capacity degradation factor and the losses due to internal resistance dissipation. 11 of 18

12 The internal resistance modelling is variable to account for (dis)charge rates as well as thermal effects (heatup during discharge and cool-down during charge). Of particular interest is the successful modelling of the transition between discharge and charge. The model is covered in more detail in Ref Model.. Telemetry Battery Voltage (V) :00 Figure 7. Battery voltage for Venus Express Deep Discharge Test 9, performed on DoY The battery voltage can be seen to drop past the critical level where a faster rate of discharge begins. The model fitting, used to assess degradation, can be seen to fit 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 sufficient to obtain a significant excursion of the EMF curve outside its linear initial part. The discharge down to 60-65% DoD showed the portion of the curve where the steeper second part started, the exact position of which is indicative of the battery degradation level. Example telemetry and model fitting of such a test is shown in figure 7. These dedicated Deep Discharge Tests were performed in visibility (for safety reasons) and were triggered by rotating the Solar Arrays edgeon to the Sun. The visibility requirement causes the discharge rate to be relatively high as the X-band transmitter is on. On-board protection was added to ensure that at a given battery voltage the arrays would rotate back to the normal position even in case of loss of ground contact. Both Mars Express and Venus Express flight control teams have determined long term trends of the capacity evolution (which is useful for long term mission planning and approval of mission extensions) and used it for short term prediction of the battery voltage behaviour for activities where the power margin available is critical. Establishing a better relation between SoC and voltage allowed the flight control team to be able to use the margin, when needed, more safely. Any allowed DoD figure in the mission planning cycle can be coupled to the corresponding voltage used and remaining capacity. 15:00 Time 16:00 12 of 18

13 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 features a wide variability in sun incidence angle on the spacecraft and periods of eclipse where no solar array power is 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 performance evaluation Provision of adequate redundancy and protection circuitry to avoid failure propagation and to ensure recovery from any malfunction within the subsystem and/or load failure. The CryoSat-2 EPS comprises the following units: A two panel fixed (non-rotating) GaAs Solar Array with 11 electrical sections per panel, regulated by means of a shunt system A single 78 Ah battery, consisting of 52 strings with 8 Li-ion cells each, manufactured by ABSL Space Products. Combined Power Control and Distribution Unit Battery charge is controlled by the PCDU in order to prevent the battery from experiencing thermal stress. During insufficient solar array power, either due to eclipse or high demand, for example in a dawndusk orbit with sub-optimal array alignment, the energy stored in the battery will be used to satisfy the bus power demand. The Battery is charged by applying an IV-method with charge current as the control parameter besides battery voltage. Eight commandable end of charge (EoC) voltage limits (steps of 200 mv) allow compensation of any potential cell parameter variations of the battery. B. Battery Usage Profile The CryoSat-2 orbit is a polar orbit but it is not sun-synchronous and consequently the orbital plane rotates with respect to the sun direction. The nodal plane regresses at a rate of about 0.25 per day. It therefore makes half a revolution, sampling all local solar times, in just over 8 months. This means that the satellite faces great variations in solar illumination and there are periods when it flies along the dawn-dusk line and is in constant sunlight, but with only one of the solar arrays illuminated. At other periods it flies in the noon-midnight plane with both solar arrays illuminated and undergoes eclipses. At maximum extent the eclipse 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 to compensate 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 considered important to take measures to reduce the rate of battery degradation in order to preserve the resource of battery life for future use. As for Mars Express, CryoSat-2 obeys the 5 key principles of Li-ion battery life preservation: 13 of 18

14 Minimise depth of discharge This is performed de facto on CryoSat-2, rather than actively managed as with Mars Express. This is thanks to the greater design margins in the CryoSat-2 power system and the more stable and repetitive nature of its power demand. This allows the operators to ensure that no excessive discharges will occur as long as routine operations are conducted. Minimise number of cycles As with Mars Express, the cycles from eclipse periods are unavoidable. However, outside of eclipse periods, 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 still mostly low enough to prevent excessive capacity fade degradation. The increased temperature (with respect to Mars Express) prevents the increase in internal resistance at lower temperatures, which has two drawbacks: Lower charge/discharge efficiency, as the internal losses are higher Lower charge capability, as the end of charge voltage limit is reached already at a lower SoC, due to 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 even when the missions DoD is around 30%. This unnecessary margin in the power system has been used on CryoSat-2 to reduce the degradation due to ageing by minimising the SoC of the battery during the whole mission. For example, there could be no energy demand from the battery for several weeks every few months when the spacecraft is in a dawn-dusk orbit. During these battery non-operational periods the SoC of the battery is reduced in order to reduce the ageing degradation. However, it is important to ensure that the reduction in SoC during mission does not interfere with the energy demand during full operation. During the transition between no eclipse to longest eclipse the battery voltage at the end of the discharge is monitored and when it crosses a predefined threshold the end of charge is modified such as to always have enough power available to ensure the safety of the satellite even in a worst case anomaly, 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 ensured by 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 the internal resistance and the battery open circuit voltage is not achievable in flight. In fact the CryoSat-2 housekeeping telemetry provides us only with the battery charge/discharge current and the battery terminal voltage, along with a summary of these in the on-board software. In addition, the solar arrays have a fixed position 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 flight data. 1. Battery internal resistance estimation One of the best ways in flight to estimate the battery internal resistance is to induce a high impulsive current demand and measure the corresponding terminal voltage drop, a method detailed in the CryoSat-2 User Manual. 3 This method works well if the current demand is large but the current demand on CryoSat-2 is typically quite low, with the largest peaks being around 3 amps, for example when a heater is switched on. However, even with this low current the method can still be applied to see if trends on the evolution of the battery internal resistance appear over satellite lifetime. 14 of 18

15 2. Battery Capacity Estimation To estimate capacity of the battery without deep discharges, the CryoSat-2 User Manual 3 proposed a method based on comparing the voltage drop and integrated current produced by the battery. Using the theoretical curve for a new battery which converts voltage into state of charge provided by ABSL, one can get the state 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 does not 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 Ah converted from the voltage drop seen in TM using the ABSL curve and the effective Ah measured by telemetry for each longest eclipse. The ABSL curve provides a conversion between voltage level of the battery and state of charge and can be used therefore to produce an estimate of the Ah that a given voltage drop should produce for a battery with a given capacity. Periodically we compare the power that should be produced by a battery with a capacity as estimated at the previous longest eclipse with the power produced now computed by 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 accurate relative value of capacity lost between tests. Using the percentage of reduction between the power produced at the previous longest eclipse and the one at the last eclipse measurement gives us a factor to apply to the last estimated capacity of the battery to estimate the current capacity of the battery. This iteration was started using the capacity given by ABSL for the battery at beginning of life. The results of this method are shown in figure 8 below and show a reasonable match with the prediction made by ABSL. Remaining Battery Capacity (%) Evaluated Capacity. from TM. ABSL Battery Capacity. Prediction 65 67% Capacity Specified EoL Performance CryoSat-2 Launch 2011 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 been evaluated using telemetry Time of 18

16 V. Summary of Li-ion Battery Management on other ESOC Missions As mentioned previously, ESA s Proba-1 mission, launched in 2001, was the first spacecraft to orbit Earth with commercial Li-ion battery technology on board. Since then almost all ESA missions have been equipped with Li-ion batteries, gradually bringing to an end the era dominated by conventional nickel-cadmium and nickel-hydrogen batteries for spacecraft. The management and usage of Li-ion batteries differs from mission to mission depending on their orbit profile and power system design. As has already been mentioned, what all missions using Li-ion batteries have in common are the important operational constraints for maximum charge and discharge voltages and currents, correct settings for end of charge levels and the avoidance of thermal stress. All this is commonly taken into account and managed by the PCDU in combination with respective thermal regulation which by design optimises the battery lifetime and reduces the operational overhead. Furthermore the level of operational experience has increased with time and missions, providing a set of applied best-practices contributing to the battery longevity, and extending missions considerably as highlighted in the above case studies. 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 and their respective battery management characteristics. While the above case studies apply to missions with power systems designed to cope with regular eclipse encounters or, in case of Mars Express, operations way beyond their nominal lifetime, other missions exist where lithium-ion batteries are not used in routine flight at all or have shown to require no additional operational measures. For example, Herschel and Planck, each equipped with a 36 Ah Li-ion battery, were flying in an orbit without eclipses where together with the solar array 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-ion batteries do not require reconditioning and are free of memory effects, no activities for capacity preservation or capacity measurements were needed throughout the mission. GOCE had a 78 Ah battery at the front of the spacecraft which operated flawlessly throughout the mission, encountering seasons of 16 eclipses per day with peak durations up to 32 minutes. Except for staying in the defined boundaries in which the battery had to be operated no particular battery management strategies had to be applied due to the mission life-limiting factor being propellant, not battery life. Battery performance assessment using in-flight telemetry to compute the battery capacity spent during discharge cycles showed good performance of the batteries. A battery simulation on ageing and number of charge/discharge cycles based on in-flight telemetry conducted by industry in preparation for the low orbit operations in 2012 concluded a battery degradation due to ageing and cycling of about 6%, whereas 19.5% was expected at that time. Shortly before the re-entry the battery was successfully operated at temperatures above 80 degc for 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 fly-by in February 2007 when the spacecraft entered solar-eclipse and for solar array performance tests in 2010 and 2014 (before and after deep space hibernation) where a step-wise sun off-pointing of the solar arrays made use of the batteries to supply the required power and assess how much power the solar arrays could provide. To preserve the battery capacity 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 a detailed understanding of the degradation level and solar eclipse seasons could only be survived with the onboard battery. To determine the battery capacity Deep Discharge tests through solar array off-pointing were performed, as with Mars Express. The collected data was fitted to dedicated models from which predictions on the capacity degradation and internal resistance trend could be made. Furthermore, the end-of-charge level was lowered outside eclipse seasons to 80%. Having kept the batteries in a good shape will be crucial in the 2014 period where the batteries will be used to allow extended periods of low aspect angles on the arrays and high DoDs during aerobraking in the Venusian atmosphere. 16 of 18

17 Table 2. Summary of Li-ion battery monitoring and management strategy for ESOC missions. Mission Battery Usage Profile Lifetime Preservation Measures GOCE Venus Express Rosetta Launch, eclipses and peak power demands Launch, eclipses and peak power demands Launch, Mars swing-by in Feb (eclipse) and Solar Array performance tests in 2010 and 2014 Capacity In-Flight Measurement None Performed In-flight telemetry used for battery simulation tests Battery end-of-charge Deep discharge tests and level lowered to 80% SoC Monte-Carlo fitting Battery end-of-charge None performed level lowered to 80% SoC 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 operators working with Lithium Ion batteries. 1. All operational lifetime preservation measures presented in this paper are useful, even though individually they may seem minor. Each measure, when applicable, allows to slow down the capacity degradation by up to a few percent over mission lifetime with respect to the spacecraft manufacturer s prediction. 2. The mission, spacecraft and power system design heavily influence the ability to perform lifetime preservation measures. Operators should work with the spacecraft designers to ensure as far as possible that the design allows the flexibility to perform these measures. 3. Operators should also work with the spacecraft designers to ensure that sufficient telemetry is available in flight 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 in order to match the real degradation. 5. The most effective (and only absolute) way to reliably measure true degradation of a Lithium Ion space battery in flight is through a Deep Discharge Test. 6. The ability to perform a Deep Discharge Test should be included in the design of spacecraft power systems, even if they normally would not have that ability (i.e. CryoSat-2). The mechanism used to do this would have to be carefully considered to ensure it could be performed at negligible risk to the mission. 7. Measures to preserve and monitor Lithium Ion batteries should be implemented by operators as part of the operations concept of a mission, planned for before launch, to maximise their effectiveness. While it is never too late to implement such measures, they are most beneficial if started from day one. 8. A close working relationship between operations teams, spacecraft and power system designers and battery experts is key to minimising the degradation and maximising the usage and lifetime of a space battery. VII. Conclusion This paper has reviewed some background on Lithium Ion space batteries and focussed on their use and management on missions flown from ESOC. Even in this small selection it is clear that the strategies being 17 of 18

18 employed are variable. There are a number of reasons for this, as highlighted in the paper. These include mission design factors, such as the usage profile of the battery, the length of the mission and the hardware available. However, the operational factors of how the battery is monitored and operated are also varied across missions. Ultimately though, it is clear that there are a number of common threads that can be followed to best monitor Lithium Ion batteries and to reduce the rate of degradation. The lessons learned point to the need to implement such strategies from the moment spacecraft operations start and to consider them in the spacecraft design and operations concept. It is also clear that there is still harmonisation to be done between different operators of Lithium Ion batteries, spacecraft designers and battery experts. This will result in an optimal realisation of the potential of this battery technology both in the spacecraft design phase and for many 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 excellent overview 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 partners that have assisted us in building up our operational experience and knowledge of working with Lithium Ion batteries. References 1 Dudley, G. and Verniolle, J., Secondary Lithium Batteries for Spacecraft, ESA Bulletin,, No. 90, May 1997, pp Sousa, 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 Astrium GmbH,., CryoSat-2 User Manual, CS-MA-DOR-SY of 18