Haa

Figure 102.- Solar array and share mode operating points for conditions near Mars.

(table 24) to the unregulated dc power bus to drive its operating point to the higher voltage stable point at point A where the array alone supported the electrical needs of the spacecraft. The circuitry consisted of redundant share mode detectors and a single boost converter that were located in the battery electronics module. The redundant share mode detector circuit sensed the voltage level of the unregulated dc bus. Should it drop to 32.6 ± 1.2 V, the share mode detector actuated nonlatching redundant battery electronic relays K2 and K3, that activated the boost converter. A timing circuit controlled the time of the relay closure to control the width of the pulse output from the boost converter at 0.5 ± 0.2 sec. The timing circuit also limits the relay closure repetition rate to 7.5 ± 2.5 sec. If the boost converter was enabled, the solar array must be Sun acquired in order to boost. The share mode detector ceased to call for boost pulses when the unregulated dc bus voltage became 34 V or higher. The boost converter circuit, powered by the battery, was a self-oscillating dc-dc converter. It provided a high voltage pulse (table 24) to the unregulated dc bus that was designed to shift the operating point at B to A in figure 102. If the VO load was too great to be supported alone by the array, the boost attempts were unsuccessful until the automatic CCS reduction of loads or VO power management by ground command diminished the total VO electrical load to a level that could be accommodated by the array. It was the backup CCS load reduction feature that removed the requirement to design a redundant boost converter circuit. The K1 relay in the battery electronics subassembly was set to enable the boost converter, and VO telemetry indicated the status of this relay. The K2 and K3 relay closures, that were sensed by the timing circuit, were also monitored to be counted by the CC.S. When the boost converter was inhibited so was the CCS boost count and load reduction feature. Logic incorporated into the share mode detector circuit automatically inhibited the circuit when the array was not Sun acquired. If the circuit was enabled by command, it must also have received the Sun-gate signal in order to function. This signal, provided by the ACS Sun sensors, indicated that the solar array was Sun acquired. The share mode detector was automatically inhibited in the absence of the Sun-gate signal. The design eliminated extra VO commanding to inhibit the circuit when boosting would be futile; these would be the occasions when planned maneuvers from the Sun caused expected share with battery discharge and the maximum power output of the array was too diminished by the turn from the Sun to support the VO load. Boosting during these occasions unnecessarily further drained the batteries. If a load decrease permitted successful boosting while off the Sun line, an artificial Sun gate would enable the share mode circuit with the issuance of a (7D) command. Attitude shifts from the Sun caused shadows to be cast upon the solar array by the low-gain antenna, array latches and hinges, and other VO structures. As can be expected, the array power output was diminished by the shadows. However, it was very difficult to accurately analyze the exact decrease of array power after a given maneuver. An adequate estimate of shadows was achieved with a study of shadow patterns cast upon the solar array of a small model of the VO. This investigation was performed in the Celestarium facility at JPL where photographs were taken of the VO model as it was oriented to collimated Sun rays at known yaw and roll angles. Output power estimates of an array that was expected to be shaded as the result of a VO maneuver included factors derived from the photographs. The photographs also yielded the derating factors to be stored as part of the PWRS computer program data base. Although shadowed array sections could generate some electrical power, a worst-case approach was used to simplify the analyses by assuming zero power output for these sections.

Tests showed the nickel-cadmium battery to be capable of reliably sustaining thousands of discharge-charge cycles when the battery was well managed. From the battery point of view, it was well managed when it operated in a cool environment and was maintained in energy balance. Battery energy balance was achieved when VO operations did not excessively drain the battery and operations insured adequate recharge before the battery was again used. A well managed battery system from the mission point of view was to have the batteries fully charged ready for any planned sequence requiring its use or ready for any unscheduled event that suddenly required battery support. It was one of the responsibilities of the power analyst to aid mission sequence design to reconcile both battery management concepts. For most mission activities, the depth of discharge of the VO batteries was limited to 45 percent, although discharge in excess of 45 percent was accommodated during Sun occultation intervals.

Battery charger lockup describes a condition of the charger when it was unable to transfer to the low charge rate mode from either its medium or high charge rate modes. Normally, such transfer was automatic when the battery voltage reached 38.25 ± 0.25 V or the battery temperature reached 29° ± Io C (85° ± 3° F). But in lockup, the charger could not transfer to its low-rate mode, automatically or by command. The lockup condition was due to an undesired logic state caused by the inability of the differential amplifier circuit in the charger to reset. The circuit had inherent hysteresis. The lockup condition was caused by a procedure that inadvertently permitted the issuance of a medium or high charge rate command to the charger after it had just automatically transferred to its low-rate mode and before the differential amplifier circuit had been permitted to reset. The charger would then charge at its newly commanded medium or high rates, but it would not automatically transfer to the low rate after the battery reached the prescribed conditions. This mode endangered the battery because of the high temperature generated by the battery overcharging in the medium or high rates. Lockup was avoided after automatic transfer to the low rate because of battery voltage limiting if the battery voltage was permitted to decline to about 36 V before the higher charge rate commands were issued to the charger. The low battery voltage reset the differential amplifier to regain the automatic low-rate transfer feature at the higher battery voltage. If the transfer was caused by high battery temperature (29° ± 1° C (85° ± 3° F) ) , the battery had to cool to 24° C (+ 3° or - 1°) (75° F (+ 5° or - 3°)) to achieve the reset. Conditions had to be present to satisfy both the battery voltage and temperature reset requirements for the proper operation of the automatic transfer to low rate. If lockup occurred, the charger was commanded off, and the open-circuit battery would then return to conditions that reset the circuit for automatic low charge rate transfer. The charger was then switched on and the unit operated normally. If the lockup went unnoticed, the battery temperature increased sufficiently to cause the 38° C (100° F) battery thermal switch to close and signal the CCS to turn off the charger.

The 10-month cruise period to Mars caused changes in the battery that resulted in temporarily diminished capacity. Continuing the battery operation in low-rate overcharge for long periods without deep discharges was thought to generate crystalline changes in the battery plates that increased the internal electrochemical impedance of the cell. The result was end-of-charge battery voltage increase and discharge voltage decrease. The impact to mission operations was that in discharge the battery voltage declined more quickly to its lower specification limit (27.4 V) before its anticipated W-hr capacity was realized. This battery condition was reversible with deep discharge and recharge. Whether such battery conditioning was required was determined from ground tests. VO batteries followed on ground the sequences undertaken by the flight batteries, and their battery voltage regulation was studied when simulated Mars orbit insertion sequence loads were applied after the long cruise period. Similar tests studied the battery response to other critical sequences. When the data suggested the need for battery conditioning before the mission event, the VO batteries would be discharged at a high rate for a time using VO loads applied to cause a deliberate share mode. This discharge sequence was followed by another having the batteries discharged with only its test loads at a safer rate until the desired depth of discharge was reached. The batteries were then recharged to complete its conditioning cycle.

Interfaces

Table 29 lists users of the various power busses, both those users that could be switched and those that could not. Note that an IRU could not be switched from its 400-Hz inverter.

The FDS provided a 4.8-kHz clock pulse signal to the power subsystem to synchronize the 2.4-kHz inverter. Magnitudes of current, voltage, and temperature sensed by PWRS sensors were indicated by equivalent voltage output from

0 0

Post a comment