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Polar mode Controller Anomaly Report

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POLAR Mode Controller Anomaly Report


An Anomaly Review Team (ART) was formed to investigate the apparent spontaneous switch of the Power Supply Electronics (PSE) Mode Controller (MC) onboard the POLAR spacecraft on September 7, 2003. The anomaly occurred while in a real-time contact following a long duration eclipse. Two parameters were observed to change state from primary to secondary values. The primary focus of this ART was on the possible reasons for the MC switch. This report documents the history and status of the MC, the investigation undertaken by the ART, and the findings of that investigation. The conclusion is that an unknown failure within the power system has caused a permanent switch to the secondary MC. This switch over will not impact the operation of the power system or the overall spacecraft operations for the remainder of the mission.


The Electrical Power Subsystem (EPS) consists of power generating (solar array), control (PSE), and storage functions (batteries). The sun’s solar energy is directly transferred from the solar arrays to the load bus. The main load bus provides regulated +28V direct current power to the instruments and spacecraft subsystems.

The PSE is the central controlling element in the EPS. It controls and maintains the +28V regulated bus by either shunting excess solar power or by boosting the power load bus voltage with stored energy from the batteries. Failure detection circuitry provides automatic switchover from primary to secondary control functions.
The MC maintains the regulated bus by controlling the charge, boost, and shunt functions within the EPS. Both primary and secondary MC circuits are active simultaneously to provide on-line redundancy. The failure detection circuitry protects against either improper bus voltage or mode overlap (e.g. shunt drives active while in boost mode). If a malfunction is detected within the primary MC, the secondary MC takes over active control and begins regulating to +28V. A simplified functional block diagram for the MC is shown in Figure 1.
Bus over and under-voltage protection is provided by automatically switching the MC and it’s associated shunt control amplifier. Over-voltage in excess of +29.5V and under-voltage of less than 27.0V causes a switchover. There was no indication in spacecraft telemetry that either of these situations occurred.
Mode overlap protection switches the MC to backup whenever the batteries are commanded by the MC to provide power to the bus (boost mode) while the shunt drive is active (shunt mode). This is the mechanism that is suspected of causing the fail-over. Switching from secondary to prime can only be accomplished by ground command.

Figure 1: Mode Controller Block Diagram

The anomaly occurred at 06:24:44 GMT on September 7, 2003 during a real-time contact. During this contact two parameters in the power subsystem were observed to change from primary to secondary values. At that time the anomaly was considered to be caused by a single event upset because all EPS parameters were nominal both before and after the event.

The first parameter (OMCSTAT) changed its value from PRIMARY (raw value of 1) to SECNDARY (0). This parameter represents the status of the PSE-MC. The internally redundant Mode Controller is responsible for transitioning the PSE seamlessly between three power modes: Charge, Boost, and Shunt.
The second parameter (OSHDRACS) changed its value from AMP_A (1) to AMP_B (0). This parameter represents the connection status of Shunt Drive A, the first of two Shunt Drives (Shunt Drive B is the other). The Shunt Drive can be connected to either Shunt Control Amplifier A (AMP_A) or Shunt Control Amplifier B (AMP_B).

Current Status

Currently the secondary MC is still active and using Shunt Control Amplifier (SCA) B. All attempts to command the MC and SCA to primary have stopped.

ART Charter

  1. Analyze all relevant telemetry and other data to assess, if possible, the status of the mode controller, and the potential for recovery of its functionality.

  2. Review all relevant telemetry to determine, if possible, the reason for the switchover.

  3. Determine whether any additional commanding of the primary MC should be done.

  4. Issue a recommendation regarding the configuration of the EPS control circuitry.

  5. Produce a report detailing the findings of the ART.

ART Membership

Scott Glubke / GSFC 444 / SSMO Spacecraft Engineering Lead (ART Chairman)

Steve Odendahl / GSFC 444 / POLAR Mission Director
Patrick Crouse / GSFC 444 / SSMO Deputy Project Manager

Anisa Ahmad / GSFC 563 / Power Systems Engineer

Steve Hearn / LMSO / GGS-POLAR Spacecraft Lead Engineer

Barbara Giles / GSFC 692 / POLAR Project Scientist

Mike Machado / HTSI / POLAR FOT Lead

ART Investigation

Telemetry review

The Flight operations Team (FOT) first noticed the anomaly when the ground system event monitor noted that two parameters in the EPS had changed state during a real time contact. At the time of the anomaly the EPS was in charge mode following a 130-minute eclipse. The solar arrays were in full sun and providing ample charge current (12.8 amps). The main bus was regulated at +28.1V. All other EPS parameters appeared normal. The FOT verified that the rest of the spacecraft subsystems and instruments were nominal. Due to the severity of the eclipse season, which lasted until September 19, 2003, no action was taken before then to reconfigure the PSE because changing the PSE requires taking the instrument high voltage off-line and thus interrupting science.

Failure Analysis

Subsequent ART meetings discussed the anomaly and various related parameters both before and after the anomaly. There were no obvious indications as to why the switchover occurred.
The GSFC Power System Engineer, Anisa Ahmad, reviewed the anomaly with other engineers within her branch to see if anyone had similar experiences on other programs. The Polar Orbiting Environmental Satellite (POES) had a similar anomaly during ground testing in February 1999. The POES spacecraft has a similar power system design and was built by the same company.
The following is a brief summary of the POES failure investigation:

The POES spacecraft has the following Mode Controller Fault Detection & Correction (MC-FDC) Circuitry:

  1. If VBUS < 27.0V for 2+/-0.5 sec, then (a) MC is changed from primary to backup, and (b) PSA-A is disconnected so both Partial Shunt Drive -1 and Partial Shunt Drive-2 are driven by PSA-B.

  1. If VBUS > 29.5V for 130 msec OR If VBUS > 30.5V for 100 msec, then

    1. MC is changed from primary to backup, and

    2. PSA-A is disconnected so both Partial Shunt Drive-1 and Partial Shunt Drive-2 are driven by PSA-B.

  1. IF MC PRI is driving BVR AND PSA-B driving shunts for >0.5 sec, then disconnect PSA-B so that Partial Shunt Drive-2 is also connected to PSA-A.

  1. If BVR is active AND BCRs are active, then

    1. MC is changed from primary to backup and

    2. PSA-A is disconnected so both Partial Shunt Drive-1 and Partial Shunt Drive-2 are driven by PSA-B.

  1. If BVR is active AND PSAs are active, then

    1. MC is changed from primary to backup and

    2. PSA-A is disconnected so both Partial Shunt Drive-1 and Partial Shunt Drive-2 are driven by PSA-B.

In summary, the two mode overlap conditions monitored by the MC-FDC are:

  1. The first is the PSA-A DRIVE being active while either the primary Mode Controller (MC) Boost or Charge modes are active. After 1.5 to 2.0 seconds, the FD commands the MC to backup, and switches the source of PSA DRIVE-1 from the PSA-A DRIVE output of the MC to the PSA-B DRIVE output.

  1. The second is the PSA-B DRIVE being active while either the Primary MC Boost or Charge modes are active. In this case, the FD switches the source of PSA DRIVE-2 from the PSA-B DRIVE output of the MC to the PSA-A DRIVE output after approximately 500 milliseconds. This detector is inhibited if the MC has been previously commanded to backup.

The POES anomaly is narrowed down to the presence of negative voltage on PSA Drive lines. The PSA drive voltages are sensed by MC-FDC to determine if the PSAs are ON or OFF. A voltage greater than 1.9 volts is interpreted as PSAs are ON. Otherwise, they are OFF. This signal is processed by an operational amplifier IC LM139. The positive rail of this circuitry is tied to about +20 volts and the negative rail is tied to spacecraft single point ground (SPG). The LM139 IC has a specified limit that any of its inverting or non-inverting terminals shall not see a voltage >300mV lower than the voltage at the negative rail. If it exceeds, then the output of this op-amp goes HIGH and is interpreted falsely as if PSAs are ON, even though the PSA drive signals is <-300mV. The MC-FDC issues commands (if PSAs are On while BVR is ON) to reconfigure the power system thinking that in fact mode controller has failed. However, in reality, the mode controller has not failed. The MC-FDC circuitry should have been designed such that it would disregard a negative voltage and not misinterpret it as a failure.

The POES anomaly investigation concluded that

  1. the electrical ground support equipment caused the POES anomaly,

  2. the worst case conditions on orbit should not cause the spacecraft to experience the MC switchover.

It was also noted that barring any other failures such a switchover would not affect the spacecraft operations and that the MC could be returned to the primary side by ground command.
Because of the science impact of cycling the high voltages, the operation to switch the MC back to the primary side was delayed until a planned maneuver in January 2004 when the instrument were planned to be off. The SCA would be left on the secondary side.

Additional Testing/Investigation

On January 12, 2004 the FOT attempted to switch the MC back to the primary side at the beginning of a planned pass. The command to switch Mode Controllers was sent from the proper procedure at 23:20:53 GMT. The command was successfully received by the spacecraft, but failed to change the state of the Mode Controller. The command was sent a second time from the procedure at 23:23:06 GMT. Again, the spacecraft acknowledged receipt of the command, but no change of state occurred. No further action was taken during that contact.
The FOT reviewed the command definition to determine if a reset, reconfiguration, or constraint removal prior to sending the MC switch command was required. No other commands or resets were identified. Because the command was received, acknowledged, and processed by the spacecraft, the MC should have switched back to the primary side. A review of the telemetry did not show even a momentary switch back to the primary MC. The MC-FDC reacts much faster than the telemetry updates so even a momentary switch may not be seen in telemetry.
The power systems group reviewed the initial data again in addition to further researching the MC design. A final suggestion was to command the shunt drives back to the primary SCA. If this command failed to switch back to primary, then it can be concluded that most probable degradations/failures that would have occurred may be:

  1. A component failure in the primary mode controller circuitry.

  1. The op-amp(s) used in the fault detection (in specific mode overlap portion) circuitry would have degraded. The op-amp used changes its output states even when its non-inverting input becomes more negative compared the op-amp’s power supply negative rail.

If either one of the above is the cause for the failure, then the FOT cannot command either mode controller or the shunt drive-A to primary sides.
On March 22, 2004 at 19:43:51 GMT, the FOT sent a command to the spacecraft to reconnect Shunt Drive A to PSE Amp A. The command was received and accepted by the spacecraft, but the configuration of the Shunt Drive did not change. No other change to the Power subsystem was observed. The command was sent a second time at 19:44:59 GMT without affect. No further commanding was attempted.

ART Discussion & Findings

The exact cause of the MC switchover cannot be determined due to the limited information available in telemetry. All attempts to switch back to the primary side have failed. The most probable location of the failure is within the fault detection circuitry. This is based on all the EPS telemetry being nominal and with FDC limits both before and after the anomaly was first noticed.

Final Disposition & Recommendation

The final recommendation of the ART is to complete the mission using the secondary power system mode controller and shunt drive amplifier. The spacecraft is currently healthy and operating properly on the secondary electronics. No other actions are necessary.

Page of Apr. 7, 04

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