HISTORY OF FLIGHTOn February 15, 2020, at about 7:42 AM pacific standard time (PST), Spirit Airlines Flight 1818, an Airbus A319-132, N521NK, experienced a fault with both engine integrated drive generators (IDG) while on approach to the Sacramento International Airport (SMF), Sacramento, California. According to the flight crew, while the airplane was descending from 4,000 feet to 2,000 feet during their approach into SMF, they heard a “clicking noise” behind the captain’s seat that sounded like an electrical component or contactor opening or closing. They also observed the “ELEC GEN 2 FAULT” message on the lower Electronic Centralized Aircraft Monitor (ECAM) and a FAULT caption illuminate on the overhead panel. Within about 11 seconds, they heard another “clicking noise” and observed the “ELEC GEN 1 FAULT” message appear on the lower ECAM. These faults resulted in the loss of electrical power supplied to both electrical networks (AC BUS 1 and AC BUS 2) leading to the loss of several flight deck displays and systems. Upon the dual loss of power, the ram air turbine (RAT) automatically extended and began driving the emergency generator to provide electrical power to vital services. Due to the nature of the emergency and the phase of flight, the flight crew declared an emergency and asked for an immediate visual approach to runway 34L. The aircraft landed without incident, was met by fire trucks, and safely towed to the gate where the passengers were deplaned normally. None of the 117 occupants aboard the airplane were injured. The regularly scheduled passenger flight was operating under the provisions of Title 14 Code of Federal Regulations Part 121 from the McCarran International Airport (LAS), Las Vegas, Nevada to SMF. AIRCRAFT INFORMATIONThe Airbus A319, in common with other Airbus aircraft, is equipped with an electronic instrument system (EIS). This system consists of six liquid crystal cockpit display units (DU), two displays in front of each pilot and two central displays. The upper display is the Engine and Warning Display (E/WD), the lower display is the Systems display. The pilot displays are part of the electronic flight instrument system (EFIS) and provide primary flight instrumentation information on the primary flight display (PFD) and the navigation display (ND). The PFDs present information on aircraft attitude, performance, flight path and autopilot modes and the NDs provide navigation, weather radar and Traffic alert and Collision Avoidance System (TCAS) information. The System Display (SD) has the capability to display 13 different system pages, the cruise page or the status page. The display has two areas: the upper section of the screen provides information based on the selection of the display, and the lower section contains permanent data that is always present regardless of the page selection. This permanent data contains information on the total and static outside air temperatures, the time, the aircraft’s gross weight and its center of gravity. In flight, the ‘default’ cruise page is generally displayed. This page shows additional engine parameters, such as fuel burn, oil quantity and vibration levels, as well as cabin air and pressurization parameters. The lower ECAM display unit normally provides the “System Display”, which presents synoptic diagrams showing the status of various aircraft systems. The ECAM display is controlled through the ECAM Control Panel (ECP), located on the center pedestal directly below the ECAM displays. A specific system page such as the “ELEC System” may be called up manually, by selection of the appropriate button on ECP; a page will appear automatically following an aircraft system failure. For both the synoptic diagrams and the control panel captions, normal system conditions are displayed in green or white and abnormal conditions in amber. A number of fault conditions also cause the red Master Warning or amber Master Caution caption lights on the flight deck to illuminate and a continuous or single chime to sound. The electrical power system (EPS) on this aircraft consists of a three-phase 115/200 V 400 Hz constant-frequency AC system and a 28 V DC system. The EPS comprises two electrical networks, a left and a right, denoted AC BUS 1 and AC BUS 2. There is also a third network, called the “AC Essential” Bus, which is supplied by either AC BUS 1 or AC BUS 2; this BUS supplies power to most of the critical aircraft systems. AC BUS 1 and AC BUS 2 networks are normally independent of one another, so that the failure of one network should not affect the operation of the other. The power supplies for flight critical systems are for the most part segregated, so that the loss of a single power source should not cause concurrent failures of systems necessary for continued safe flight. Two engine driven AC generators (GEN 1 & GEN 2), one mounted on each engine, normally power the EPS system. Each generator is driven from the engine high-pressure spool via an engine accessory gearbox and an integrated hydro-mechanical speed regulator. The regulator transforms variable engine rotational speed into a constant-speed drive for the generator. The constant-speed drive and the generator together form an assembly known as an Integrated Drive Generator (IDG). According to Spirit Airlines records, the Number 1 IDG (IDG1) was installed on the aircraft on November 12, 2017, in Philadelphia, PA. It had been installed for approximately 827 days and accumulated 8,582 hours and 3,693 cycles prior to its removal. The Number 2 IDG (IDG2) was installed on the incident aircraft on March 29, 2017, in Orlando, FL. It had been installed for approximately 1,055 days and accumulated 10,584 hours and 4,581 cycles prior to its removal. The output frequency and voltage of each IDG is controlled by its respective generator control units (GCU); the Number 1 GCU controls the IDG1 and the Number 2 GCU controls the IDG2. The GCUs also protect the network by controlling the associated generator line contactor (GLC). The EPS also comprises a third generator (APU GEN) that is driven directly by the Auxiliary Power Unit; it produces the same electrical output as each of the main engine generators. Additionally, a ground power connector near the nosewheel allows ground power to be supplied to all busbars. A Ground and Auxiliary Power Control Unit (GAPCU) regulates the frequency and voltage of the APU generator; it also protects the network by controlling the external power contactor and the APU generator line contactor. In the event there is a loss of both the AC BUS 1 and AC BUS 2 busbars in flight, vital aircraft electrical services can be supplied by an Emergency Generator which is driven by the ram air turbine (RAT). The RAT deploys either automatically, usually because of loss of both main AC busbars, or on manual selection. RAT deployment is indicated by a green icon on the ECAM hydraulic system page. The electrical system is controlled via the electrical panel located on the overhead console in the cockpit. This panel provides for annunciation of the status of the electrical system and fault conditions. The generator control switches are identified as GEN 1, GEN 2 and APU GEN for the left, right and APU generators. If a fault occurs with a generator, an amber fault message will illuminate on the respective generator control push button (PB) switch indicating the channel is offline. For a generator fault, pressing the PB once turns the generator to OFF. Pressing the PB a second time is required to reset the generator back to ON. A review of maintenance records found that N521NK had experienced a loss of power supplied from IDG1 and IDG2 prior to the February 15th, 2020, flight to SMF. The following paragraphs describe the four previous events. According to Airbus, the introduction section of their troubleshooting manual (TSM) indicates that if the operator cannot find a fault symptom and/or fault isolation procedure necessary to ensure the continued airworthiness of the aircraft, or if the operator thinks that the information given is not complete, they should contact Airbus. On January 19, 2020, N521NK, operating as Spirit Airlines Flight NK1848, experienced a loss of all electrical power and the loss of nose wheel steering about 3 minutes after shutting down the Number 1 engine while taxing after landing at the Orlando International Airport (MCO), Orlando, Florida. Spirit Airlines Maintenance performed a run-up of the Number 1 engine per the procedures contained in aircraft maintenance manual (AMM) 71-00-00. The full electrical load was supported by IDG1, and no faults were noted. The aircraft was determined to be ok for continued service and no additional troubleshooting was accomplished. According to Airbus, they were not contacted regarding this event. On January 23, 2020, N521NK, operating as Spirit Airlines Flight NK1445, experienced a loss of IDG1, nose wheel steering and all electrical power during a single engine taxi after landing at MCO. The flight crew observed the ‘ELEC GEN 1 FAULT’ message on the lower ECAM display. Spirit Airlines Maintenance performed an operational test of IDG1 per AMM 24-22-00. No faults were noted. According to Airbus, Using the Airbus AirnavX Troubleshooting tool, TSM task 24-00, “GEN 1 FAULT Warning” could have been used. This task requires the operator to perform an engine run and verify the GEN1 parameters. If the fault is confirmed (i.e. GEN1 parameters are not correct), then the troubleshooting data (TSD) from the GCUs shall be extracted. If the fault cannot be confirmed, Airbus can be contacted. According to the NVM download from GCU Number 1, N521NK experienced a loss of IDG1 while performing a single engine taxi after landing on January 29, 2020. No maintenance information was available from Spirit regarding this event. On February 14, 2020, N521NK, operating as Spirit Airlines Flight NK806, experienced a loss of all electronics (dual generator failure) while on approach to the Hartsfield-Jackson Atlanta airport (KALT), Atlanta, Georgia. Spirit Airlines Maintenance performed a check of the oil levels on the Number 1 and 2 engine IDG and found the levels good per aircraft maintenance manual (AMM) 12-13-24. Maintenance also ran both engines per AMM 71-00-00 and found that they both checked good. The aircraft was determined to be ok for continued service. According to Airbus, the Airbus AirnavX Troubleshooting tool, TSM task 24-21-51, “IDG1” and “IDG2” could have been used. These TSM procedures require checking the wiring of the IDG servo-valve, and then changing the GCU if the wiring is correct, and ultimately replacing the IDG if the fault continues. Collins Aerospace performed a download of the non-volatile memory (NVM) from the Number 1 and Number 2 GCU’s and the GAPCU at their facility located in Rockford Illinois. The data revealed that during the February 14 and 15th flights in which there was a dual loss of electrical power, both GCU’s recorded fault code 145. According to Collins, GCU’s provide electrical system monitoring and protection and will automatically turn off their respective electric power generating channel when design parameters are exceeded. Each IDG contains an electro-mechanical servo valve which plays a role in frequency control. GCU’s monitor servo valve frequency modulation and if it detects that the frequency is outside of specified design limits, it will trigger fault code 145. This fault is designed to stop system operation in the case of a failing servo valve. For the February 14th and 15th flights, the Number 2 GCU tripped at a normal load as fault 145 and the Number 1 GCU tripped as fault 145 after the electrical load was doubled due to the Number 2 generator dropping off-line. The NVM data also revealed that the Number 1 GCU recorded fault code 145 on January 19, 23, and 29th 2020. For these events, Collins noted that fault code 145 occurred after the airplane had landed and was performing a single-engine taxi. In this configuration, the Number 1 IDG is supplying all of the aircraft electrical load. Collins Aerospace also performed examination and testing of the Number 1 & 2 IDG’s removed from the incident airplane. Testing found that both IDGs failed their frequency control tests. The frequency test data for IDG1 showed a frequency modulation that would have triggered fault 145 on its GCU. IDG2 showed similar frequency issues. However, despite extensive testing by Collins and Airbus with varied electrical load and acceleration/decelerations profiles, it was not found that IDG2 produced the exact frequency modulation to specifically trigger fault 145. It was noted that the test setup was not identical to the on-aircraft situation. Upon disassembly of IDG1, it was observed that the fixed hydraulic cylinder blocks had been modified by a non-OEM process and were marked as PN CB-31773. Visual inspection of the fixed blocks revealed that portions of the Number 3 and 5 bronze bore lining were missing, having flaked off after wearing thin. There was approximately 1/4 to 1/3 of the bore missing in the Number 3 and Number 5. Several “flecks” of metal, copper in color were noted in the block assembly. The incident fixed cylinder block and its pistons from IDG1 were replaced by a new set for the purpose of re-testing the frequency behavior of the IDG. During the re-test, it was observed that the IDG produced the proper frequency performance and also fully passed the Acceptance Test Procedure (ATP). The teardown of IDG2 was generally unremarkable. Like IDG1, the fixed cylinder block had been modified by the same FAA-approved minor repair and marked as PN CB-31773. The fixed-block showed severe wear in several bores. Again, visual inspection revealed that portions of the bronze bore lining were missing, having flaked off after wearing thin. Because the “SV MODULATION,” fault 145 indicates a fault with the servo valve, it was removed from IDG2 for isolated testing. The servo valve passed all tests and no faults were found. After replacing only the damaged fixed-block/pistons, IDG2 passed ATP. All frequency performance anomalies noted before the repair were remedied by replacing the fixed block. For thoroughness, both IDGs were disassembled completely in order to inspect for any additional contributing factors. None were found. During the teardown of IDG1, to confirm the non-contribution of the servo valve, it also was removed for isolated testing, which it passed. The fixed cylinder blocks and mating piston assemblies from each IDG were submitted to the NTSB Materials Laboratory for examination. The fixed cylinder blocks were identified as part number 764635; the cylinder block removed from the left engine, S/N 1266, was vibro-peen marked with S/N Y10-245-19 and the cylinder block removed from the right engine, S/N 1157, was vibro-peen marked with S/N Y04-232-49. As previously noted, both fixed cylinder blocks were also vibro-peen marked with ‘CB-31773’, which was reported as the process number for rework performed on the cylinder block bore linings by an aftermarket company. The part numbers and serial numbers listed above were assigned by the original equipment manufacturer (OEM) of the fixed cylinder blocks. The fixed cylinder block S/N Y10-245-19 (left engine) and associated pistons were visually examined and photo documented. Bands of circumferential wear marks were visible on the outer diameter (OD) of the pistons. A few pistons had heavier bands of wear than others, and the bands occurred along most of the length of the pistons. Wear was observed in the gold-colored liners assembled inside the bores of the fixed cylinder block. The wear was uneven around the circumference and occurred close to either end of the bores, with the heaviest wear extending up from the bottom of the bores. The gold-colored liners were worn through in some areas in several bores, and silver-colored base metal was visible. Areas on bores 3, 5, and 6 were missing segments of gold-colored liner; the areas with missing liner ranged in size from very small to almost half the liner length. In areas where the gold-colored liner was beginning to wear through, thin circumferential rings became visible. In areas where the gold-colored liner was missing, thicker circumferential rings consistent with machining were visible. A piece of gold-colored debris was present in the bottom of at least one of the bores. For additional information, reference the Materials laboratory factual report for this incident. The fixed cylinder block Y04-232-49 (right engine) and associated pistons were visually examined and photo documented. Only bore 3 had the gold-colored liner worn through. The heavily worn bore 3 and another bore with a lesser amount of wear (bore 7) were destructively examined. The bottom 0.1 inches of the sectioned piece of liner was missing, and the edge of the liner adjacent to the missing portion had thinned. The ID surface of the block piece and OD surface of the liner had a substance on it consistent with the bores examined from the other fixed block. The substance along the bottom half of the liner OD had darkened in color, along with several areas on the ID of the block piece. Debris was visible adhering to the substance inside some of the machined grooves in the ID of the block. For additional information, reference the Materials laboratory factual report for this incident. AIRPORT INFORMATIONThe Airbus A319, in common with other Airbus aircraft, is equipped with an electronic instrument system (EIS). This system consists of six liquid crystal cockpit display units (DU), two displays in front of each pilot and two central displays. The upper display is the Engine and Warning Display (E/WD), the lower display is the Systems display. The pilot displays are part of the electronic flight instrument system (EFIS) and provide primary flight instrumentation information on the primary flight display (PFD) and the navigation display (ND). The PFDs present information on aircraft attitude, performance, flight path and autopilot modes and the NDs provide navigation, weather radar and Traffic alert and Collision Avoidance System (TCAS) information. The System Display (SD) has the capability to display 13 different system pages, the cruise page or the status page. The display has two areas: the upper section of the screen provides information based on the selection of the display, and the lower section contains permanent data that is always present regardless of the page selection. This permanent data contains information on the total and static outside air temperatures, the time, the aircraft’s gross weight and its center of gravity. In flight, the ‘default’ cruise page is generally displayed. This page shows additional engine parameters, such as fuel burn, oil quantity and vibration levels, as well as cabin air and pressurization parameters. The lower ECAM display unit normally provides the “System Display”, which presents synoptic diagrams showing the status of various aircraft systems. The ECAM display is controlled through the ECAM Control Panel (ECP), located on the center pedestal directly below the ECAM displays. A specific system page such as the “ELEC System” may be called up manually, by selection of the appropriate button on ECP; a page will appear automatically following an aircraft system failure. For both the synoptic diagrams and the control panel captions, normal system conditions are displayed in green or white and abnormal conditions in amber. A number of fault conditions also cause the red Master Warning or amber Master Caution caption lights on the flight deck to illuminate and a continuous or single chime to sound. The electrical power system (EPS) on this aircraft consists of a three-phase 115/200 V 400 Hz constant-frequency AC system and a 28 V DC system. The EPS comprises two electrical networks, a left and a right, denoted AC BUS 1 and AC BUS 2. There is also a third network, called the “AC Essential” Bus, which is supplied by either AC BUS 1 or AC BUS 2; this BUS supplies power to most of the critical aircraft systems. AC BUS 1 and AC BUS 2 networks are normally independent of one another, so that the failure of one network should not affect the operation of the other. The power supplies for flight critical systems are for the most part segregated, so that the loss of a single power source should not cause concurrent failures of systems necessary for continued safe flight. Two engine driven AC generators (GEN 1 & GEN 2), one mounted on each engine, normally power the EPS system. Each generator is driven from the engine high-pressure spool via an engine accessory gearbox and an integrated hydro-mechanical speed regulator. The regulator transforms variable engine rotational speed into a constant-speed drive for the generator. The constant-speed drive and the generator together form an assembly known as an Integrated Drive Generator (IDG). According to Spirit Airlines records, the Number 1 IDG (IDG1) was installed on the aircraft on November 12, 2017, in Philadelphia, PA. It had been installed for approximately 827 days and accumulated 8,582 hours and 3,693 cycles prior to its removal. The Number 2 IDG (IDG2) was installed on the incident aircraft on March 29, 2017, in Orlando, FL. It had been installed for approximately 1,055 days and accumulated 10,584 hours and 4,581 cycles prior to its removal. The output frequency and voltage of each IDG is controlled by its respective generator control units (GCU); the Number 1 GCU controls the IDG1 and the Number 2 GCU controls the IDG2. The GCUs also protect the network by controlling the associated generator line contactor (GLC). The EPS also comprises a third generator (APU GEN) that is driven directly by the Auxiliary Power Unit; it produces the same electrical output as each of the main engine generators. Additionally, a ground power connector near the nosewheel allows ground power to be supplied to all busbars. A Ground and Auxiliary Power Control Unit (GAPCU) regulates the frequency and voltage of the APU generator; it also protects the network by controlling the external power contactor and the APU generator line contactor. In the event there is a loss of both the AC BUS 1 and AC BUS 2 busbars in flight, vital aircraft electrical services can be supplied by an Emergency Generator which is driven by the ram air turbine (RAT). The RAT deploys either automatically, usually because of loss of both main AC busbars, or on manual selection. RAT deployment is indicated by a green icon on the ECAM hydraulic system page. The electrical system is controlled via the electrical panel located on the overhead console in the cockpit. This panel provides for annunciation of the status of the electrical system and fault conditions. The generator control switches are identified as GEN 1, GEN 2 and APU GEN for the left, right and APU generators. If a fault occurs with a generator, an amber fault message will illuminate on the respective generator control push button (PB) switch indicating the channel is offline. For a generator fault, pressing the PB once turns the generator to OFF. Pressing the PB a second time is required to reset the generator back to ON. A review of maintenance records found that N521NK had experienced a loss of power supplied from IDG1 and IDG2 prior to the February 15th, 2020, flight to SMF. The following paragraphs describe the four previous events. According to Airbus, the introduction section of their troubleshooting manual (TSM) indicates that if the operator cannot find a fault symptom and/or fault isolation procedure necessary to ensure the continued airworthiness of the aircraft, or if the operator thinks that the information given is not complete, they should contact Airbus. On January 19, 2020, N521NK, operating as Spirit Airlines Flight NK1848, experienced a loss of all electrical power and the loss of nose wheel steering about 3 minutes after shutting down the Number 1 engine while taxing after landing at the Orlando International Airport (MCO), Orlando, Florida. Spirit Airlines Maintenance performed a run-up of the Number 1 engine per the procedures contained in aircraft maintenance manual (AMM) 71-00-00. The full electrical load was supported by IDG1, and no faults were noted. The aircraft was determined to be ok for continued service and no additional troubleshooting was accomplished. According to Airbus, they were not contacted regarding this event. On January 23, 2020, N521NK, operating as Spirit Airlines Flight NK1445, experienced a loss of IDG1, nose wheel steering and all electrical power during a single engine taxi after landing at MCO. The flight crew observed the ‘ELEC GEN 1 FAULT’ message on the lower ECAM display. Spirit Airlines Maintenance performed an operational test of IDG1 per AMM 24-22-00. No faults were noted. According to Airbus, Using the Airbus AirnavX Troubleshooting tool, TSM task 24-00, “GEN 1 FAULT Warning” could have been used. This task requires the operator to perform an engine run and verify the GEN1 parameters. If the fault is confirmed (i.e. GEN1 parameters are not correct), then the troubleshooting data (TSD) from the GCUs shall be extracted. If the fault cannot be confirmed, Airbus can be contacted. According to the NVM download from GCU Number 1, N521NK experienced a loss of IDG1 while performing a single engine taxi after landing on January 29, 2020. No maintenance information was available from Spirit regarding this event. On February 14, 2020, N521NK, operating as Spirit Airlines Flight NK806, experienced a loss of all electronics (dual generator failure) while on approach to the Hartsfield-Jackson Atlanta airport (KALT), Atlanta, Georgia. Spirit Airlines Maintenance performed a check of the oil levels on the Number 1 and 2 engine IDG and found the levels good per aircraft maintenance manual (AMM) 12-13-24. Maintenance also ran both engines per AMM 71-00-00 and found that they both checked good. The aircraft was determined to be ok for continued service. According to Airbus, the Airbus AirnavX Troubleshooting tool, TSM task 24-21-51, “IDG1” and “IDG2” could have been used. These TSM procedures require checking the wiring of the IDG servo-valve, and then changing the GCU if the wiring is correct, and ultimately replacing the IDG if the fault continues. Collins Aerospace performed a download of the non-volatile memory (NVM) from the Number 1 and Number 2 GCU’s and the GAPCU at their facility located in Rockford Illinois. The data revealed that during the February 14 and 15th flights in which there was a dual loss of electrical power, both GCU’s recorded fault code 145. According to Collins, GCU’s provide electrical system monitoring and protection and will automatically turn off their respective electric power generating channel when design parameters are exceeded. Each IDG contains an electro-mechanical servo valve which plays a role in frequency control. GCU’s monitor servo valve frequency modulation and if it detects that the frequency is outside of specified design limits, it will trigger fault code 145. This fault is designed to stop system operation in the case of a failing servo valve. For the February 14th and 15th flights, the Number 2 GCU tripped at a normal load as fault 145 and the Number 1 GCU tripped as fault 145 after the electrical load was doubled due to the Number 2 generator dropping off-line. The NVM data also revealed that the Number 1 GCU recorded fault code 145 on January 19, 23, and 29th 2020. For these events, Collins noted that fault code 145 occurred after the airplane had landed and was performing a single-engine taxi. In this configuration, the Number 1 IDG is supplying all of the aircraft electrical load. Collins Aerospace also performed examination and testing of the Number 1 & 2 IDG’s removed from the incident airplane. Testing found that both IDGs failed their frequency control tests. The frequency test data for IDG1 showed a frequency modulation that would have triggered fault 145 on its GCU. IDG2 showed similar frequency issues. However, despite extensive testing by Collins and Airbus with varied electrical load and acceleration/decelerations profiles, it was not found that IDG2 produced the exact frequency modulation to specifically trigger fault 145. It was noted that the test setup was not identical to the on-aircraft situation. Upon disassembly of IDG1, it was observed that the fixed hydraulic cylinder blocks had been modified by a non-OEM process and were marked as PN CB-31773. Visual inspection of the fixed blocks revealed that portions of the Number 3 and 5 bronze bore lining were missing, having flaked off after wearing thin. There was approximately 1/4 to 1/3 of the bore missing in the Number 3 and Number 5. Several “flecks” of metal, copper in color were noted in the block assembly. The incident fixed cylinder block and its pistons from IDG1 were replaced by a new set for the purpose of re-testing the frequency behavior of the IDG. During the re-test, it was observed that the IDG produced the proper frequency performance and also fully passed the Acceptance Test Procedure (ATP). The teardown of IDG2 was generally unremarkable. Like IDG1, the fixed cylinder block had been modified by the same FAA-approved minor repair and marked as PN CB-31773. The fixed-block showed severe wear in several bores. Again, visual inspection revealed that portions of the bronze bore lining were missing, having flaked off after wearing thin. Because the “SV MODULATION,” fault 145 indicates a fault with the servo valve, it was removed from IDG2 for isolated testing. The servo valve passed all tests and no faults were found. After replacing only the damaged fixed-block/pistons, IDG2 passed ATP. All frequency performance anomalies noted before the repair were remedied by replacing the fixed block. For thoroughness, both IDGs were disassembled completely in order to inspect for any additional contributing factors. None were found. During the teardown of IDG1, to confirm the non-contribution of the servo valve, it also was removed for isolated testing, which it passed. The fixed cylinder blocks and mating piston assemblies from each IDG were submitted to the NTSB Materials Laboratory for examination. The fixed cylinder blocks were identified as part number 764635; the cylinder block removed from the left engine, S/N 1266, was vibro-peen marked with S/N Y10-245-19 and the cylinder block removed from the right engine, S/N 1157, was vibro-peen marked with S/N Y04-232-49. As previously noted, both fixed cylinder blocks were also vibro-peen marked with ‘CB-31773’, which was reported as the process number for rework performed on the cylinder block bore linings by an aftermarket company. The part numbers and serial numbers listed above were assigned by the original equipment manufacturer (OEM) of the fixed cylinder blocks. The fixed cylinder block S/N Y10-245-19 (left engine) and associated pistons were visually examined and photo documented. Bands of circumferential wear marks were visible on the outer diameter (OD) of the pistons. A few pistons had heavier bands of wear than others, and the bands occurred along most of the length of the pistons. Wear was observed in the gold-colored liners assembled inside the bores of the fixed cylinder block. The wear was uneven around the circumference and occurred close to either end of the bores, with the heaviest wear extending up from the bottom of the bores. The gold-colored liners were worn through in some areas in several bores, and silver-colored base metal was visible. Areas on bores 3, 5, and 6 were missing segments of gold-colored liner; the areas with missing liner ranged in size from very small to almost half the liner length. In areas where the gold-colored liner was beginning to wear through, thin circumferential rings became visible. In areas where the gold-colored liner was missing, thicker circumferential rings consistent with machining were visible. A piece of gold-colored debris was present in the bottom of at least one of the bores. For additional information, reference the Materials laboratory factual report for this incident. The fixed cylinder block Y04-232-49 (right engine) and associated pistons were visually examined and photo documented. Only bore 3 had the gold-colored liner worn through. The heavily worn bore 3 and another bore with a lesser amount of wear (bore 7) were destructively examined. The bottom 0.1 inches of the sectioned piece of liner was missing, and the edge of the liner adjacent to the missing portion had thinned. The ID surface of the block piece and OD surface of the liner had a substance on it consistent with the bores examined from the other fixed block. The substance along the bottom half of the liner OD had darkened in color, along with several areas on the ID of the block piece. Debris was visible adhering to the substance inside some of the machined grooves in the ID of the block. For additional information, reference the Materials laboratory factual report for this incident. FLIGHT RECORDERSThe flight data recorder was a Honeywell 4700 solid-state flight data recorder (SSFDR) that records airplane flight information in a digital format using solid-state flash memory as the recording medium. The FDR was examined upon receipt, was found to be in good condition and the data were extracted normally from the recorder. The FDR recording contained approximately 27 hours of data. Two flight segments on the recording were of interest to the investigation: the last flight (Flight 1818) and flight NK806 which occurred two segments before the last flight on the recording. An uneventful flight, flight number NK805, was performed between the two legs of interest. The cockpit voice recorder (CVR) was a Honeywell 6022 solid-state CVR that records at minimum two hours of digital audio from four input channels – one channel for each crew member (captain, first officer, and observer) and one channel for the cockpit area microphone (CAM). The four channels are processed to create five distinct audio files: one 30-minute file for each channel, one 2-hour file for the CAM, and one 2-hour file of channels 1, 2, and 3 combined. Upon arrival at the laboratory, it was evident that the CVR had not sustained any heat or structural damage, and the audio information was extracted from the recorder normally, without difficulty. The recording began with a crew discussing non-pertinent events in cruise and ended after the aircraft had landed and parked at a gate at KLAS. It was determined that the recording was not of the event flight (NK1818), but rather the flight prior to the event flight (NK805). Further, it was determined that the aircraft had a similar inflight dual generator failure the previous day on flight NK806, which was two flight legs prior to the event leg. The operator was notified that the CVR did not contain audio from the incident flight. The operator stated that a request was made from their maintenance and engineering department to pull the CVR circuit breaker on the aircraft after the dual generator failure on flight NK806 to preserve the recording from that flight. However, the request to pull the circuit breaker in KATL did not reach the appropriate maintenance personnel before the aircraft departed for KLAS. The CVR circuit breaker was then pulled after the aircraft arrived in KLAS. As a result, the CVR audio preserved was from the flight KATL to KLAS resulting in the CVR not being powered during the incident flight. ORGANIZATIONAL AND MANAGEMENT INFORMATIONSeveral improvement actions were taken by Airbus to reduce the likelihood of additional events in which there is a loss of all electronics resulting from a dual generator failure. Airbus has improved their A318/A319/A320/A321 trouble shooting manual (TSM) by incorporating steps to direct maintenance towards a direct extraction of the post flight report (PFR) and troubleshooting data (TSD) from the GCUs. Airbus is also reviewing the possibility of introducing new recurrent maintenance tasks in their maintenance planning document (MPD) to ensure the IDG’s are operational. The objective would be to detect a degraded IDGs/APU GEN and to replace them before they fail in service.