Battery Thermal Design Conception of Turkish Satellite

Murat Bulut1, Selman Demirel,2 Nedim Sozbir3 and Senol Gulgonul4

Turksat International Satellite and Cable Operator, Golbası, Ankara, 06839, Turkey

This paper describes some general aspects of the thermal design of Turkish satellite and presents a detailed analysis of the battery and its compartment. The battery compartment is analyzed, using mathematical modeling, to calculate the radiator area and heater power consumption. The battery internal temperature gradients are estimated, also using mathematical modeling, to verify the cells package design. The thermal package was used to create the thermal model where the temperature distributions are predicted by a thermal analyzer.

Nomenclature AIT = assembly integration and test facility BCRB = battery connection relay box BMS = battery management software BOL = beginning of life CM = communication module EOL = end of life EPC = electrical power conditioner EQ = equinox EPS = electrical power subsystem GEO = geostationary K = conduction conductance, W/K Li-ion = lithium-ion Mc = thermal mass, J/K MIMU = miniature inertial measurement unit MLI = multi-layer insulation OSR = optical solar reflector PCU = power conditioning unit PDR = preliminary design review PFLDIU = platform and payload distribution unit Q = heat rate, W q = heat rate, W/m2 S/C = spacecraft SM = service module SMU = satellite management unit SS = summer solstice STR = star tracker T = temperature, K or oC t = time, s TCS = thermal control subsystem TMM = thermal mathematical model T/O = transfer orbit WS = winter solstice α = absorptivity ε = emissivity

ijℑ = radiation exchange factor (script F) between surfaces (or nodes) i and j

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1 Satellite Thermal Specialist, Satellite Design and R&D Center, Konya yolu 40.Km Golbasi, Ankara. 2 Satellite Electrical Power Specialist, Satellite Design and R&D Center, Konya yolu 40,km Golbasi,Ankara. 3 Consultant, Assistant Professor, Satellite Design and R&D Center, Konya yolu 40.Km Golbasi, Ankara. 4 Satellite Design and R&D Center Director, Satellite Design and R&D Center, Konya yolu 40.Km Golbasi, Ankara

6th International Energy Conversion Engineering Conference (IECEC)28 - 30 July 2008, Cleveland, Ohio

AIAA 2008-5787

Copyright © 2008 by Nedim Sozbir. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

Subscripts i = location; node; inner i,j = from node, body, or surface i to node, body, or surface j j = location; node; inner j,k = from node, body, or surface j to node, body, or surface k Superscripts A = associated with albedo d = associated with dissipated heat E = associated with emitted Earth radiation S,s = associated with solar radiation r = associated with radition

I. Introduction ithium-ion (Li-ion) batteries are excellent candidates to provide power and energy storage for satellites in GEO due to their high specific energy, high energy density, and excellent low temperature performance. Therefore,

Turkish satellite uses Li-ion batteries. The satellite electrical power subsystem (EPS) consists of two kinds of power supply. As a primary power supply that uses solar arrays and as a secondary power supply that uses Li-ion battery .As a secondary power supply, the energy which stored in the batteries maintain the continuous power for the peak and eclipse power demands. Li-ion batteries are the last generation type of batteries which is mainly prefer due to high energy capacity and the low heat dissipation. Batteries are the one of the most important equipment for the GEO satellites thermal design because of their high dissipation values and narrow operating temperature range. Batteries have special requirements that must be taken into consideration during satellite thermal design. Temperature and temperature uniformity are important factors for obtaining optimum performance from the satellite battery module. Li-ion battery specification require an optimum operating temperature around 20 oC and battery thermal design should ensure the optimum operating temperature and temperature uniformity during all the mission of the spacecraft from beginning of life (BOL) to end of life (EOL).

L

II. Battery Design Concept The energy needs of the satellite supply by electrical power subsystem. In the satellite EPS, as one of the main

power supplier, two module Li-ion batteries are used. The Li-ion battery is used when solar array power is not sufficient during pre-launch operation (ground power is not available), launch, transfer orbit (T/O), eclipses, peak discharge during sunlight. These battery modules are mounted on the north and south panels in the service module(SM).

At the battery design all the calculation is made according to the worst case principle. The main orbital phase is eclipse phase that the battery is used as a single power supplier. According to the power budget, the worst case satellite power consumption is calculated for the eclipse phase. The battery sizing and design is made to compensate the satellite power consumption during the longest eclipse period. The longest eclipse duration is 72 minute and that means that the designed battery gives all the satellite power need include peak values during 72 minute.

The battery cell capacity value is the most important unit at the battery design. This value is different from manufacture to manufacture and that is obvious that, high cell capacity means low battery mass. Nowadays Li-ion battery technology offers the highest cell capacity per kg and hence this technology commonly preferred at the space application. Such as all batteries, some losses occur at the Li-ion battery. For the battery sizing, these losses must be taken into account. There are four kinds of losses which are battery harness losses, discharge current losses, calendar losses, cycling losses due to the different phenomenon. The available energy is a function of the leakage current of the cells. This loss is also reversible. The charge/discharge equinox cycles induce a loss called cycling losses which is irreversible. Cycling loss is 2% of energy loss for a 15 years mission (1380 cycles). The aging of the 15 years of the mission induce a loss called calendar loss which is irreversible too. Calendar loss is 3% of energy loss for a 15 years mission. Aging and cycling are considered as independent and additive. Charge-discharge cycling also causes some losses in the battery. According to the battery's cell capacities and satellites needs, with taking account the losses, the number of cells that used in battery is calculated. Li-ion battery has serial and package design. Cells are connected in parallel and

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create the package form (2p-3p-4p). With the serial connection of the package, the serial design is formed. In design two 3p-12s Li-ion battery module are used. This means that in one module there is 3 parallel and 12 serial cell (total 36 cell in one module). The Li-ion batteries are full autonomous. The battery management software (BMS) is implemented in satellite management unit and operates the battery without ground control. The battery management is performed by software providing 2 high level functions. The charge management function offers an automatic way to control the battery without ground intervention. The monitoring and failure detection function allows the autonomy in case of failures. That means the battery management not only offers an automatic way to control the battery without ground intervention, but also ensures the autonomy in case of failures. At the battery design the battery cell protection also must be taken into account. The protection of the satellite battery cell open circuit failure is made by a network of by-pass device .This device is implemented with each cell package (one per cell package). One of the most important topics for Li-ion battery is thermal regulation. In Li-ion battery the thermal dissipation is very low compare to the other kinds of space batteries. In spite of the low dissipation values, the thermal management of the Li-ion batteries is very critical. As an example for Li-ion battery, the continuous temperature during equinox phase of GEO life time is from +15 to 30 oC. This values shows that the thermal design of the battery directly affects the battery design and besides the efficiency of the battery. The heaters are implemented on every cell for thermal regulation. These heaters are switched on /off with the help of thermal regulation software. This software is also full autonomous and can operate without ground intervention

III. Battery Thermal Control Design Description Li-ion batteries are mounted inside spacecraft and they are located on north and south SM panels, as shown in Figure 1. The thermal control design of each battery module is shown in Figure 1. Consequently, the batteries

thermally insulated from the internal part of the spacecraft and have their own thermal control. Thermal design for batteries includes a multi-layer insulation (MLI) blanket cover, and an optical solar reflector (OSR) is provided on the honeycomb panel to radiate the heat out from the batteries1. The battery radiator size was 0.132 m2. MLI is used to isolate the batteries from the satellite interior. Batteries are fixed to a honeycomb wall fitted with OSR. The optical property values along with the thermophysical properties of the material used are shown in Table 1.

Figure 1. Schematic of the battery location

The battery thermal designs use radiators and controlled heaters via thermistors. The radiator is sized to keep the batteries below the maximum allowable temperature under worst hot-case conditions and controlled heaters are used to maintain minimum allowable temperatures under cold-case conditions. The temperature range of batteries is different from other SM units.

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Cells

MLI

HoneycombBaseplateOSR

Heaters on each cell

Interfiller

Figure 2. Battery thermal design

Table 1. Thermophysical and optical properties of various elements. Components Solar Solar Emissivity Thermal Conductivity absorptivity absorptivity Conductance

(BOL) (EOL) W/m2K W/mK OSR 0.11 0.266 0.84 External MLI 0.35 0.5 0.61 External BLACK MLI 0.93 0.93 0.84 Internal MLI N/A N/A 0.05 Baseplate-honeycomb 3200 Lateral 3.5 Transversal 1.35

For passive thermal control, thermal interfiller is used to ensure good and removable thermal contact with radiator. Radiative insulation which is MLI is used to decouple batteries from the internal thermal environment of the satellite. In order to keep the temperature gradient with narrow range, dedicated OSR radiators are mounted under batteries module.

For active thermal control, heaters and thermistors are used. Heater lines are implemented on every battery module. Heaters located on the upper part of the baseplate. 1.53 W heater is located per cell. Thermistors are located and fixed on the baseplates of the coldest packages.

IV. Battery Thermal Model External heat fluxes incident on each element of the satellite depend on position of satellite in orbit, satellite

attitude in its current position, relative positions of the sun and earth, eclipse time, and satellite geometry2. The external fluxes considered for the calculations are 1371 W/m2 during equinox (EQ), 1418 W/m2 during winter solstice (WS), and 1326 W/m2 during summer solstice (SS).

Knowledge of the external heat fluxes and the dissipation of the elements allow the temperature distribution variations to be determined as function of thermal couplings, area, solar absorptivity, and emmisivity. These parameters are controlled, and combinations are selected to produce the desired temperature for a given dissipation.

The governing equation is energy equation that consists of transient, conduction, and radiation terms plus boundary conditions (solar, albedo, earth radiation) as a source term. This equation can be written as

)()()()( 44jki

jijjri

ri

jiji

EEAAssssdi

ii TTKTTAqAqAqAQ

dtdT

Mc −−−ℑ−+++= ∑∑ σσεαα (1)

In the above equation, the term on the left-hand side shows the thermal capacitance of the element. The first term

on the right-hand side refers to operational load, the second term is the net external radiant flux absorbed, and the third term in the same side is the net radiation rejected, whereas the last term is the internal conduction3.

V. Thermal Analysis of Batteries The battery thermal model includes an estimate of the temperature changes during charging/discharging of Li-ion cell in the thermal environment which affects the battery charge/discharge characteristics. Batteries have wide variations in orbital dissipation and show significant transients despite their relatively large mass because of repeated charging and discharging4. Thermal dissipation varies with state of charge, temperature, and charge rate, and it may differ for batteries of the same general type. The thermal conditions of satellite batteries are frequently moderated by automatic switch offs and switch ons and adjustments of recharge rates are preset temperatures. In addition, charge limits and voltage/temperature (V/T) charge control by ampere-hour integration may be regulated by onboard computers5. But management of the net heat output is normally accomplished by mounting to a thermally controlled platform4. Various ranges have been specified for Li-Ion batteries with the intention of increasing life. Batteries need to be maintained within the temperature range from 0 to 40 oC. The range +15 to 30 oC is considered standard during equinox phases of GEO lifetime. Critical applications require 20 oC during solstices phases of GEO lifetime. High temperatures are avoided by sizing the radiator to accommodate hot case conditions while the battery is in extended discharge. Heaters, activated by the thermistors, would maintain the temperature above the lower limit during a no-charge or trickle-charge operation under cold case conditions4. The battery compartment is analyzed using mathematical modeling in order to calculate the radiator area and heater power consumption. The total dissipation evolution during 72 minutes discharge for one cell is 3.33 W. Therefore, the total dissipation for each panel is approximately 120 W. The most extreme configuration is considered when the battery thermal analyses are performed. The extreme configuration is obtained by a combination of cell dissipation (discharge/charge, shunt, and cell failure), solar illumination (seasons and solar array fluxes), thermo-optical properties (BOL/EOL) and external environment (webs/reflectors).

Table 2. Cell temperatures NORTH PANEL SOUTH PANEL Cell Cell Case Condition Min Max Min Max WS EOL 100% drive HC 19.4 26.8 15.2 22.3 SS EOL 100% drive HC 11.5 20.1 20 26.8 WS EOL 100% drive HC shunt 19.4 26.9 15.2 22.3 SS EOL 100% drive HC shunt 11.6 20 19.9 26.9 EQ EOL 100% drive HC 19.6 31.1 19.9 30.1 EQ EOL 100% drive HC 1 stack failed 19.6 32 19.9 31.1 EQ EOL 100% drive HC 2 stacks failed 19.6 33.1 19.9 32.2 WS BOL 100% drive CC 19.6 26.3 18.6 25.2 SS BOL 100% drive CC 17.3 22.9 20 26.6

The thermal analyses was performed that the performance of the Li-ion battery was fulfilled with comfortable margins under worst case conditions during the whole lifetime. The qualification temperatures range of the batteries is from 0 oC to 54 oC. The extreme guaranteed temperatures were obtained from 5.4 oC to 38.2 oC at worst hot and cold cases during the whole lifetime. Cell temperatures are shown in Table 2. The temperature battery module did not exceed its qualification temperature range.

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References 1Suresha, S., Gupta, S.C., and Katti, R.A., “Thermal Sensitivity Analysis of Spacecraft Battery,” Journal of Spacecraft and

Rockets, Vol. 34, No. 3, 1997, pp. 384-390. 2Gilmore, D. G., Spacecraft Thermal Control Handbook Volume I: Fundamental Technologies, 2nd ed., The Aerospace Press,

California, 2002. 3Megahed, A., and El-Dib, A., “Thermal design and Analysis for Battery Module for a Remote Sensing Satellite,” Journal of

Spacecraft and Rockets, Vol. 44, No. 4, 2007, pp. 920-926. 4Karam, R. D., Satellite Thermal Control for Systems Engineers, Vol.181, AIAA, Reston, VA, 1998. 5Chetty, P.R.K., Satellite technology and Its applications, 2nd ed., TAB Professional and Reference Books, Blue Ridge

Summit, PA, 1991, pp.97-171.