Performance of Nickel-Cadmium Batteries on the Goes I-K Series of Weather Satellites

PERFORMANCE OF NICKEL-CADMIUM BATTERIES ON THE GOES I-K SERIES OF WEATHER SATELLITES

Sat P. Singhal Computer Sciences Corporation

Walter G. Alsbach Kris Engineering, Inc.

Gopalakrishna M. Rap NASA/Goddard Space Flight Center

Abstract

The US National Oceanic and Atmospheric Administration (NOAA) operates the Geostationary Operational Environmental Satellite (GOES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the GOES series consists of 5 spacecraft (originally named GOES I-M), three of which are in orbit and two more in development. Each of five spacecraft carry two Nickel-Cadmium batteries, with batteries designed and manufactured by Space Systems Loral (SS/L) and cells manufactured by Gates Aerospace Batteries (sold to SAFT in 1993). The battery, which consists of 28 cells with a 12 Ah capacity, provides the spacecraft power needs during the ascent phase and during the semi-annual eclipse seasons lasting for approximately 45 days each. The maximum duration eclipses are 72 minutes long which result in a 60 percent depth of discharge (DOD) of the batteries. This paper provides a description of the batteries, reconditioning setup, DOD profile during a typical eclipse season, and flight performance from the 3 launched spacecraft (now GOES 8, 9, and 10) in orbit.

National Aeronautics and Space INTRODUCTION Administration (NASA) is responsible for procurement, launch, and checkout of these spacecraft before turning them over to The US National Oceanic and Atmospheric NOAA for operational use. They were built Administration (NOAA) operates the by Space Systems Loral (SS/L) with Geostationary Operational Environmental components from many vendors. Satellite (GOES) spacecraft (among others) to support weather forecasting, severe storm A key dement in overall mission success is tracking, and meteorological research by the the successful operation of the battery National Weather Service (NWS). The latest system. These spacecraft carry two Nickel- in the GOES series consists of 5 spacecraft Cadmium batteries designed and (originally named GOES I-M), three of manufactured by SS/L with cells which (GOES 8, 9, and 10) are now in orbit. manufactured by Gates Aerospace Batteries (soldto SAFTin 1993). Thebatteries, control electronics. Key features of the PCE whichconsistof 28 cellswith a 12 Ah are as follows: capacity each, provides the spacecraft power needs during the ascent phase and during the a. Provides functions required for the semi-annual eclipse seasons lasting for primary bus. approximately 45 days. The maximum b. Provides direct energy transfer of SA duration eclipses are 72 minutes long which power to the power distribution bus. result in a 60 percent depth of discharge C. Regulates and limits the voltage of (DOD) of the batteries. the primary power bus to 42 + 0.5 V de during sunlight operation.

This paper provides a description of the d. Minimizes primary bus ripple and batteries including the design and voltage transients by use of solar manufacturing process through acceptance array regulation and primary bus testing and pre-launch preparation filtering.

(References 1-3). Results are also provided e. Allows flexibility of battery-charge from life-cycle testing at the Naval Surface control to optimize battery energy Warfare Center, Crane (Reference 4) using balance, thermal control. battery packs prepared from cells out of the f. Maintains minimum battery batch manufactured for spacecraft use. On temperature control through use of orbit data is provided from ascent phase, thermistors and heaters in the reconditioning, and from the eclipse seasons associated batteries. for the three spacecraft in orbit. g Includes provisions for individual battery reconditioning. SPACECRAFT h. Provides fail-safe and redundant actuation control of spacecraft The GOES 8, 9 and 10 spacecraft are three- Electro-explosive device (EED) (pyrotechnics) functions. axis stabilized geostationary satellites. Their mass is 999 kg and the nominal main bus Eliminates single-part failure power is 1150 W. They carry two main criticality by use of circuit redundancy, protective functions, scientific instruments (Imager and Sounder) and a number of other measuring devices. fault protection of spacecraft heater During sunlight they are powered by a single loads, and alternate mode operation wing, two panel solar array with 1300 W selectable by command.

j° Permits operational flexibility and BOL Equinox and 1050 W EOL Summer Solstice output, _during eclipse two 12 Ah status monitoring of key subsystem NiCd batteries with 28 cell each sustain the parameters by command and telemetry functions. spacecraft. Power transfer from the batteries to the spacecraft bus is achieved through Redundancy Provisions. Power diode coupling. The power control unit conditioning functions use both active and (PCU) is the principal element for commandable redundancy. Commandable management and control of spacecraft redundancy is provided for the following primary power. The PCU, one sequential PCE functions: shunt unit, and four electro-explosive device a. Battery discharge control extension units form the primary power b. Battery temperature control C. Voltage telemetry monitors mounted external to the power control unit d. Battery relay commanding provide the reconditioning load. e. Battery charge control f. Electro-explosive device (EED) Commanded Load Control. Application of ignition control (3 levels) power to individual loads is through an g. Battery reconditioning control on/off control input to each load dc/dc converter and by direct power bus switching Some important functions for the battery of non-electronic loads (heaters). control are listed below Eclipse Load Controller. The GOES PCU Battery Charge Control. The battery contains redundant load controller functions. charge control configuration provides Each load controller monitors the solar array capability of up to eight discrete battery current and the shunt current. Each primary charge rates. Battery charging current is power bus load (excepting command supplied by six charge control arrays located functions) is connected to and disconnected on the solar array wing. Within the PCU, from the primary bus by command. In charge control array sections A, B, and C are addition, power is automatically removed connected through relays and isolating from sunlight loads upon eclipse entrance diodes to battery 1. A similar arrangement is and automatically restored upon eclipse exit. provided for charge control arrays D, E, and Command override of this automatic F for battery 2. function is provided.

Battery Temperature Control. The battery BATTERY DESIGN temperature control circuitry provides dual DESCRIPTION modes to connect and disconnect the heater. A total of four battery temperature controls The battery subsystem of the GOES are provided, one for each half of each spacecraft consists of two assemblies, each battery. Automatic control of minimum containing 28 series-connected cells with a battery temperature at 5 (2 for GOES 8) +1 nominal capacity of 12 ampere-hours (Ah) °C is provided by heaters integrated with and weighing approximately 12.9 kg each. each battery assembly and separate Battery cell design has been guided by the temperature controllers in the power control life and reliability requirements of a 5-year unit. A precision thermistor on each battery geo-synchronous satellite application and provides temperature feedback to these incorporates features that reduce the effects controllers. Full override allows the heaters of unavoidable Ni-Cd cell degradation to be switched on or offby command. modes. The GOES 12 Ah battery is shown in Figure I. Battery Excessive Discharge Protection. Automatic battery excessive discharge protection is provided by sensing the voltage across seven cell groups of four cells each.

Battery Reconditioning. Individual reconditioning of each battery is provided by command. A parallel group of resistors gl

I .__ restraint of two rows of 14 cells between Electrical Design. The nominal capacity of two endplates, with each group of four cells each battery assembly is 12 Ah. The 60% supported by a rib structure. The key depth-of-discharge (DOD) design limit for component to the battery design is the maximum duration eclipse operation concept of supporting four cells on one rib, represents load capabilities of 409 W at and compressing seven of these beginning of life (BOL) and 398 W at 5 subassemblies between two endplate/tie-rod years, based on average cell discharge assemblies. The design of the rib positively voltages of 1.245 and 1.22 V, respectively, holds the corners of the prismatic cells in and a 1.0 V loss in the PCU. The 28 cells in place. The compression force exerted by the series provide compatibility with charge endplates on the cells is controlled by the voltage limits and deliver typical discharge torque applied by the two tierods. The cells voltages of 32.4 to 36.0 V. are pre-loaded to 40 lbf/in 2 in this manner.

r The end plates are designed to withstand a Reconditioning. Reconditioning of the nominal cell overcharge pressure of 75 lbffin batteries during the one-month period 2 with minimal deflection. The endplates also preceding each eclipse season is required to provide mounting flanges for supporting the maximize subsequent voltage performance two connectors. The endplates are machined and consequently minimize depth of from heat-treated 7075-T73 aluminum, discharge. which alleviates stress corrosion issues, and the tierods are titanium. The cell support Circuit Reliability. Circuit reliability within ribs are cast from A357-T61 aluminum. the battery is achieved by redundancy in wiring. Series connection between cells is Thermal Design. The thermal design of the provided by two parallel-connected, stranded GOES battery is such that temperature copper wires soldered to the cell terminal control is both passive and active. The lugs. Battery power connections are made passive control is primarily conductive to the terminals at the end of the cell series through the cell support ribs, and the active, string by four redundant wires leading to the control is via resistive heaters mounted on battery power connector. the same ribs. The seven aluminum ribs are sized to effectively conduct the heat Telemetry. The battery wiring harness dissipated by the four adjacent battery cells design includes provisions for telemetry of to the equipment platform of the spacecraft. 28 individual cell voltages and 2 battery The design is such that temperature gradients temperatures. Additionally, overall battery are only 3 °C. The ribs also have a mounting voltage and current signals can be sensed. flange for the resistive heater elements. These flanges are centrally located such that The structural/mechanical design of the an even distribution of heat throughout the GOES battery is optimized to efficiently battery assembly is achieved at times when perform two important functions: the heaters are activated. The seven heaters a. Cell support and restraint will be sized to dissipate the required heat b. Battery to equipment platform mounting for thermal control. Temperature sensing for each battery is accomplished by four Cell Support and Restraint. The basic precision wafer thermistors located on the mechanical design approach involves the battery assembly. Battery flatness is maintained within 0.020 inch during assembly to allow for uniform thermal Interfaces. The battery interfaces are rigidly contact with the spacecraft. A silicone controlled during battery assembly. The thermal grease is used between the battery battery mounting surface and mounting hole and equipment platform on the GOES pattern are controlled within close tolerances spacecraft to further reduce gradients and for an assembly of this type. Flatness of the contact resistance. battery is maintained within 0.020 inch, and the mounting hole pattern is drilled within Cell Design Summary. The nickel- 0.020 inch. Two electrical connectors are cadmium cell is designed to maximum provided on the battery assembly. The margins for a greater than 5-year orbital power connector is an eight-socket lifetime. The prismatic Ni-Cd cells are connector through which the battery is packaged in a thin wall (0.012 in) 304L charged and discharged. Four bus wires stainless steel container. A low profile each are connected to the positive and terminal seal/cover configuration is used with negative terminals. The second connector is double ceramic alumina-to-metal seals. a 50-socket connector. Through this Eleven positive and 12 negative plates are connector the interface is provided for used. The specified negative-to-positive individual cells voltage sensing, thermistor electrochemical capacity ratio is 1.70:1. This measurement, and heater bus. ratio ensures that the cells will remain positive limited during the 5-year mission. Battery and Cell l:llistory. The following The positive electrode group has a designations refer to the flight batteries. theoretical electrochemical capacity of 15.6 _+1.2 Ah and the negative group 26.4 _+_3.6 Spacecraft Battery Serial Number Ah. Active material loading is 12.5:1_-0.6 and GOES-I 204, 203 15.50 _+0.65 g/dm for the positive and GOES-J 205, 206 negative plates, respectively. The negative GOES-K 207, 208 plates are treated with Teflon to reduce cadmium migration and increase the quantity Life-test. As with most NASA spacecraft, of electrolyte used within the cell. The life-testing was started at Crane Naval balance of positive and negative loading, Surface Warfare Center. A pack of cells for electrochemical capacity ratio, and negative each satellite is undergoing real time or plate treatment enhances long-term cell accelerated life-test. The test conditions are performance required for the 5-year GOES close to those in orbit with a 6.0 A, mission. Non-woven PeUon 2536 nylon equivalent to C/2, discharge rate and 0.9 A filament material approximately 0.007-0.009 charge rate followed by a trickle charge inch thick is used for the separator. Weight period. Before entering a shadow period a concentration of 31% potassium hydroxide reconditioning cycle is run. The cycles electrolyte provides for the transport of ions shown are (Figures 2-4): between the electrodes. This combination is being used for sealed nickel-cadmium cells GOES I. Shadow Periods 2,12, and 25 ; operated over the temperature range of 0 to GOES J. Shadow Periods 2,9, and 25 ; 30 °C. Carbonate and nitrate levels are GOES K. Shadow Periods 1 and 10; maintained at less than 2.0 g/1 and 1.0 rag/l, respectively.

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Figure 4. Battery DOD and Average Cell Voltages for Battery Pack GOESK during Shadow Periods 1 and 10. GOES I 5/7/89 - 8/16/93 All curves show nominal behavior and are (GATES, FL) within specification. There is also good agreement with the recorded flight data. #3 Capacity at 20 °C: Please note that the shadow period 25 U < 1.479 V, 11.5 Ah - 12.7 Ah corresponds to 12.5 years of mission, which #4 Capacity at 10 °C: is more than twice the specified lifetime. U < 1.525 V, 12.0 Ah - 13.1Ah #5 Capacity at 0 °C: Manufacturing. The cells used for this U < 1.530 V, 12.1 Ah - 13.1 Ah battery were manufactured by GATES, FL #6 Conditioning at 20 °C from(x_ to yyy, designated as lot # 5 with U< 1.469 V, 11.8 Ah - 13.2 Ah serial numbers 12AB31 and 12AB35. Based #7 Voltage Recovery on an original 9 Ah design, they were name- U < 1.190 V plated as 12 Ah cells with specified acceptance capacity at 12.6 Ah. Out of the GOES J 5/7/91- 8/16/93 500 originally manufactured cells 20 were (GATES,FL) taken into a 500 cycle test to address problems in two Aerospace Alerts. These # 3 Capacity at 20 °C: alerts, one from 1985 reported problems of U < 1.479 V, 11.5 Ah - 12.7 Ah negative electrode failure and the other from # 4 Capacity at 10 °C: 1988 reported problems in the hot gas sinter U < 1.525 V, 12.0 Ah - 13.1 Ah process and with the Pellon 2536 separator. # 5 Capacity at 0 °C: U< 1.530 V, 12.1Ah - 13.1Ah Storage. The cells were put in storage in a # 6 Conditioning at 20 °C controlled environment U < 1.469 V, 11.8 Ah - 13.2 Ah # 7 Voltage Recovery Activation at Vendor. The basic activation U< 1.190 V consists of seven steps, some of which have been repeated until required results were GOES K 5/7/93 - 8/16/93 achieved. In the listing below, these steps (SAFT, France) are listed in their normal sequence and with their specified key values. # 3 Capacity at 20 °C: U < 1.485 V, 11.9 Ah - 12.9 All # 1 Setup # 4 Capacity at 10 °C: # 2 Letdown 0.2 Ohms U< 1.525 V, 12.0 Ah - 13.1Ah # 3 Capacity at 20 °C 1.51 V max, # 5 Capacity at 0 °C: 12.8 Ah U < 1.530 V, 12.1 Ah - 13.1 Ah # 4 Capacity at 10 °C 1.52 V max, # 6 Conditioning at 20 °C 12.0 Ah U < 1.469 V, 11.8 Ah - 13.2 Ah # 5 Capacity at 0 °C 1.54 V max, # 7 Voltage Recovery 11.5 Ah U< 1.190 V # 6 Conditioning at 20 °C 1.51 V max, 12.8 Ah Acceptance at SS/L. The basic acceptance # 7 Voltage Recovery consists of seven steps, some of which have been repeated until required results were

l0 achieved. In the listing below, these steps a° Battery voltage monitors are listed in their normal sequence and with b. Individual cell voltage monitors their specified key values. For the individual C. Battery charge current monitors batteries, the measured capacity at 20 °C is d. Battery discharge current monitors reported. e. Battery temperature monitors f Battery heater status monitors #1 Reconditioning g. Battery relay status monitors #2 Voltage Recovery h. Battery reverse current monitors #3 Capacity at 20 °C 1.51 V max, i. Battery relay control commands 12.8 Ah j. Battery charge rate control #5 Capacity at 0 °C 1.54 V max, commands

II.5 Ah k. Battery reconditioning controls #6 Conditioning at 20 °C 1.51 V max, 1. Battery charge voltage limiting 12.8 Ah m. Thermostatically controlled battery #7 Voltage Recovery heaters

Pre-launch and Launch. During the pre- GOES I 12/6/93 - 5/11/94 launch and launch phases of the mission, the #3 Capacity at 20 °C: 5A spacecraf_ is kept on external power until U < 1.489 V, 12.0 Ah - 13.0 Ah approximately 4 minutes before launch. Should a launch hold of more than 10 GOES J 12/6/93 - 5/11/94 minutes be encountered while the spacecrat_ #3 Capacity at 20 °C: 5A is on internal power, consideration should be U < 1.489 V, 12.0 All - 13.0 All given to recharging the batteries before continuing with the launch operation. GOES K 12/6/93 - 5/11/94 # 3 Capacity at 20 °C 5A Battery Charge Control. Batteries can be U < 1.484 V, 10.9 Ah - 12.9 Ah charged in either a continuous or sequenced mode. The primary consideration on the following recommended battery charge BATTERY OPERATION control implementation is to minimize battery stress while providing adequate charge The two 12 ampere-hour rated nickel- return to ensure energy balance. This cadmium batteries connected to the primary objective is satisfied using the lowest bus via redundant diodes and relays are practical charge rate in the sequenced mode designed for operation at a 60*/, maximum to minimize average battery temperature. DOD for the mission. Criteria for battery The methodology recommended has been management including power subsystem flight proven on previous SS/L programs and design features to implement the same are verified by life test cycling. discussed below. General Configuration Control. The Functions provided for each battery by the electrical power subsystem permits power subsystem to monitor, control, and considerable flexibility in battery charge protect the batteries are as follows: management. Charge power is derived from the solar array using two identical groups of

11 threechargecontrol arrays. Group 1 margin, a mission goal for maximum DOD consists of arrays A, B, and C; group 2 of should not exceed 60%. arrays D, E, and F. The typical current output of these arrays is summarized in Normal Eclipsed Orbits Charge Control. Table 1 for various conditions. Battery charge control performance is evaluated by use of the following telemetry Table 1. Battery Charge Array Current provided for each battery:

k a, Battery voltage Charge Control b. Individual cell voltages C. Charge current Arra Module d. Battery temperature Season/Life e. Battery heater status

Vernal Equinox The aplied 110% charge return at the full BOL __ rate, followed by trickle charging until the EOL next eclipse, ensures sufficient recharge for Autumnal Equinox all eclipse cycles. The rates shown above are BOL sufficient to maintain full battery charge. EOL Summer Solstice Battery Temperature Control. Battery BOL temperature profiles are largely governed by EOL the battery charge profiles. Thermostatically Winter Solstice controlled 19.5 watt battery heaters are provided to maintain the minimum battery temperature above +4 (1 for GOES 8) °C. These heaters are designed to cycle on at +5 Two basic charge modes are available. First, (2) :t:1 °C and off.at +9 (5) +1 °C. with the normal recommended synchronous Heater operation is controlled by a precision orbit sequenced mode, current is applied to thermistor in each battery in conjunction the two batteries in an alternating fashion in with level detectors and relay drivers within an approximately 10-minute cycle. Any the PCU. This function can be overridden by combination of the six charge arrays can thus command to either turn the battery heater on be switched between the batteries. or off regardless of battery temperatures. Alternatively, any combination of arrays A, The intended use of this override function is B, and C can be used for battery 1 and any to provide manual heater control in the event combination of arrays D, E, and F for battery of a failure in the automatic heater control. 2. Second is the continuous charge mode. In this mode only the latter arrangement is Battery Reconditioning. To maximize practical. battery discharge voltage during each eclipse, both batteries are reconditioned prior For one time only, a maximum DOD of 70% to the start of each eclipse season. These would be allowed for a battery temperature reconditioning cycles are performed as close maintained between 5 °C and 25 °C. This as possible to the next solar eclipse to obtain maximum DOD, if approached, allows for no the maximum benefit. margin for contingencies. To provide such a

12 ON-ORBIT BATTERY For all cases when the batteries experienced PERFORMANCE a non-zero discharge current, data was retrieved from the archives at 5.12 second

First of the GOES series I-M spacecraft was intervals. The charge removed during launched on April 13, 1994 and was renamed discharge (D) from the batteries was then GOES 8 after achieving nominal operational calculated as an integral of the discharge current over time, i.e., orbit. The next two were launched on May 23, 1995 (GOES 9) and April 25, 1997 D (Ah) = I I dt. (GOES 10). The batteries provide spacecraft power needs from just before launch to the time of partial Solar Array deployment soon The integral is evaluated numerically by after launch (the outer panel is deployed using current values at 5.12 see intervals. approximately 10 days later providing full The depth of discharge (DOD) is then given power). The batteries also shoulder the as a percent fraction of the nameplate power needs during eclipses, any maneuvers capacity (Co) of the batteries (12 Ah each causing loss of solar power, and whenever for a total of 24 Ah for the system). Thus the power produce d falls short of the spacecraft needs. The sections below discuss DOD = 100 * D (Ah) / Co battery performance during these various phases. Specific data is generally provided gives the DOD value for the individual for GOES 10. However, data from GOES 8 battery or the system as a whole depending and GOES 9 is also included whenever this on the values of the current and capacity used above. data was useful to establish trends.

Ascent phase Battery Performance Table 2 lists all times during the ascent phase when the batteries on the GOES-K

The batteries on the GOES spacecraft spacecraft were subjected to discharge. The provide power to the primary bus starting a maximum depth of discharge (DOD) was few minutes before launch when the recorded during the Mag boom and SA spacecraft is switched to internal power. deployment phase. The battery discharge This first round of battery support is currents for this specific case of are shown in completed when the Solar Array (SA) is Figure 5. partially deployed approximately 1.5 hours after launch. During the next few days, the batteries provide the spacecraft power needs during special events such as the magnetometer boom and full SA deployment phase and the dipole estimation phase. Also depending upon the time in orbit before the Apogee Maneuver Firing number 1 (AMF #1 ), the batteries provide support during several eclipses lasting approximately 15 minutes each.

13 Table 2. GOES-10 Battery Discharge History from Launch to Early On-orbit

Max Dis I Max Date DOY Time Range Event Batte_ 1 BaUer), 2 DOD (°/e) 4/25/97 115 05:43 - 07:08 Launch to Partial SA deploy 4.0 4.1 20.6

4/25/97 115 18:33 - 18:47 Transfer Orbit Eclipse #1 4.4 5.2 9.5

4/26/97 116 07:11 - 07:25 Transfer Orbit Eclipse #2 5.4 5.9 10.2

4/26/97 116 19:48 - 20:03 Transfer Orbit Eclipse #3 5.7 6.0 12.7

4/29/97 119 16:50 - 17:30" Apogee Maneuver Firing #2 1.6 1.3 4.0

5/04/97 124 23:14 - 00:01 Mag Boom & SA Deploy 4.5 4.7 26.0

5/05/97 125 18:03 - 18:30 Dipole Estimation 7.1 7.4 22.7

* Intermittent discharge currents (upto 0.6 A) from 17:30 to 18:00.

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V the battery voltage changes. Since the battery is being discharged by connecting it Battery Reconditioning to a constant resistor (139.6 Ohms), the discharge current is given by the use of SpacecraR Batteries are reconditioned prior Ohm's law to the start of each eclipse season. The batteries are individually reconditioned by I = V/R use of the following sequence after verifying that the other battery is connected to the and the charge removed as an integral of spacecraft bus. battery voltage, i.e.,

a. Command the charge to the battery to off D (Ah) = (l/R) .f V dt b. Command the battery discharge relay number 2 off In addition, since the reconditioning process c. Inhibit the battery undeervoltage continues for 60 - 65 hours, the voltage data protection for integration is sampled at 1 minute d. Command the battery recondition on intervals at the beginning and end of the process (where the voltage is changing The 139.6 ohm resistive load will be comparatively rapidly) and at 5 minute connected to the battery, resulting in an intervals during the middle 48 hour period. initial C/48 (0.25 A) reconditioning discharge rate. The individual cell voltages of Figure 7 shows the performance of GOES-8 the selected battery are monitored battery 1 during its first reconditioning cycle throughout the reconditioning discharge flail 1994). The reconditioning was period. When the first cell voltage reaches terminated when cell 12 voltage dropped to a 0.5 + 0.1 V, the reconditioning discharge is value of 0.5 V. Corresponding data for terminated. Figure 6 shows the battery battery 2 is shown in Figure 8. reconditioning circuitry.

Table 3 compares the results from all On-orbit reconditioning has been performed reconditioning efforts to date for all three prior to 7 eclipse seasons for GOES-8, and 5 GOES spacecrat_s (7 for GOES-8, 5 for eclipse seasons for GOES-9. The batteries GOES-9, and none for GOES-10). The data on GOES-10 were not reconditioned prior to show that the battery capacity has improved the fall 1997 eclipse season due to problems with time and reaches a maximum value at associated with the Solar Array anomaly approximately 130 percent of nameplate encountered after launch and early on-orbit capacity. The table also shows the end of operations. discharge (EOD) battery voltage for each case. Charge removed from the batteries during reconditioning was calculated using a different approach than that described earlier. During reconditioning, the nominal discharge current has a value of 0.25 A (C/48 rate), but the step size for the discharge current telemetry is 0.06 A, too coarse to show discharge current changes as

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N Q. central 7 days of the eclipse season, this 20- minute rule was further modified to a 10- Battery Performance During The minute rule. In addition, the AOCS team Eclipse Seasons kept both Earth Sensors operating during the eclipse season except for the days when the The GOES spacecraft experience the loss of power engineer required them to be turned solar power during the semi-annual eclipse offto avoid excessive drain on the batteries. seasons that last approximately 45 days each centered on the Vernal and Autumnal Figure 10 shows the total DOD for the equinox. The durations of the eclipses vary battery system for all 3 spacecraft for the Fall from a few minutes to a maximum of 72 1997 season. The data for GOES-8 and -9 minutes on or near the equinox. We have also show the effects of the power analyzed the eclipse data for all 13 eclipse management by modifying the times for the seasons (7 for GOES-8, 5 for GOES-9, and uplink carders. However, other aspects of one for GOES-10) and found it to be active power management are evident only consistent and predictable. We provide for the GOES-10 spacecratt since this was details of the latest (Fall 1997) eclipse season the first eclipse season for it. below along with some statistical data from the previous cases. Figure 11 shows the battery discharge currents for Day of Year (DOY) 266 Figure 9 shows the daily battery DOD during (September 23, 1997) for the 3 spacecratt. the Fall 1997 eclipse season for GOES-10. The eclipse times are separated by 2-hour The figure shows the DOD separately for intervals reflecting the fact that the 3 batteries 1 and 2 and for the battery system spacecratt are geostationary at 75 degrees W as a whole. As the eclipse duration gets (GOES-8), 105 deg W (GOES- 10), and 135 longer, the batteries are discharged for a deg W (GOES-9) longitude, The figures longer period resuting in almost a linear show a sharp rise in the discharge currents increase in the battery DOD. The DOD near the end of the eclipse due to the uplink curve, as shown, is not smooth however carriers (approximately 90 Watt increase in because of the active load management that power consumption). had to be performed to stay within the 60 percent DOD limit. Operationally, we have Figures 12 through 14 show the battery the so-called "20-minute" rule that was minimum voltage versus DOD for batteries imposed after failure of a power amplifier on 1 and 2 for GOES-8, -9, and -10 GOES-8. The uplink carders from the respectively. GOES-10 data shows larger ground stattion to the spacecraft provide spread between the battery voltages at the extra heating to the power amplifiers during same DOD (encounterd before and after the the eclipse but result in a higher DOD that maximum duration eclipse), probably would exceed the 60% limit during the reflecting the fact that these batteries had not longer eclipses. A balance is struck between gone through reconditioning, as mentioned the two requirements by keeping the carders earlier. Figure 15 shows the battery minimum up during the eclipse except during the versus DOD data for all 6 batteries together. central portion of the season when the The results clearly indicate that the batteries carders are brought down at the start of the are behaving as a family. eclipse and then brought up 20 minutes before the end of the eclipse. During the

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(A) e§ellOA mnm!u!w _Ja.e8 to 12 C for GOES-9, and 2 C to 13 C for During the Fall 1997 eclipse season, the GOES-10. Daily peaks in the battery minimum battery and cell voltages recorded temperatures were observed near the end of for the 3 spacecraft were as follows. the discharge cycle while the minimums were recorded a few hours before the end of GOES-10 I GOES-9 1 GOES-8 charge cycle, as expected. Battery 1 32.9 1 32.5 I 32.5 Table 4 lists the minimum battery voltages Battery 2 during all of the eclipse seasons encountered 32.9 [32.5 I 32.5 by the 3 spacecraft. As expected, the minimum battery voltages get lower with the Battery 1 Cells aging of the batteries but they seem to have 1.174 [1.161 1 1.160 leveled offat 32.5 Volts. The battery Battery 2 Cells temperatures have stayed in the same range 1.186 1.161 for all eclipses (values as indicated above for the latest eclipse season). The battery temperatures ranged from approximately 0 C to 9 C for GOES-8, I C

Table 4. Minimum Battery Voltages During the Eclipse Seasons

GOES-8 GOES-9 GOES-10 Battery 1 Battery 2 Battery 1 Battery 2 Battery 1 Battery 2 Fall 1994 33.3 33.3 Spring 1995 33.1 33.1

Fall 1995 32.7 32.7 33.1 33.1

Spring 1996 32.5 32.5 32.9 32.7

Fall 1996 32.7 32.5 32.7 32.7

Spring 1997 32.5 32.5 32.7 32.7

Fall 1997 32.5 32.5 32.5 32.5 32.9 32.9

specifications and results behave as a family. SUMMARY AND CONCLUSIONS In addition, the battery capacities to date have shown improvement and tend to indicate a leveling off at approximately 16 On-orbit performance of the batteries on the Ah under the C/48 discharge conditions. Life three spacecraf_ GOES 8, 9, and 10 indicates cycle tests at Crane tend to indicate that the that the batteries are performing within batteries will have no trouble providing

29 supportfor the required mission life of 5-7 years.

REFERENCES

1. GOES I-M Databook, DRL 101-08, GSFC Specification S-480-21 A, SS/L, Palo Alto, CA. 2. GOES Program, Spacecraft Operations Handbook (SOIl), Volume III (Rev G, August 1995), Spacecraft Description, SS/L, Palo Alto, CA. 3. GOES Program Presentations such as Battery CDR. 4. Private Communication. See also: Review of Ni-Cd Tests at Crane, Harry Brown and Steve Hall, Naval Surface Warfare Center, Crane; and Gopal Rap, NASA Goddard Space Flight Center.