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21 Sep 2018 - Final report released into A6-FDS 737-800 Tyre Failure and Loss of Hyd A 8 Aug 2017

Synopsis

On 8 August 2017, a flydubai Boeing B737-800, A6-FDS, operated a scheduled passenger flight FZ081, from Dubai International Airport (OMDB) to Bahrain International Airport (OBBI). There were a total of 35 persons on-board the Aircraft: 29 passengers, two flight crewmembers, and four cabin crewmembers.

The Aircraft lined up on runway 12R for departure, and after take-off clearance was given, the flight crewmembers commenced the takeoff.

After the Aircraft lifted off the runway, the landing gear was retracted, and at about 250 feet radio altitude while climbing, the Aircraft experienced a pressure loss of Hydraulic System A. The Aircraft continued climbing to the pre-selected altitude of 4,000 feet.

The flight crew contacted ATC Departure declaring that the Aircraft had suffered an hydraulic system problem, and requested to go to the hold for troubleshooting. The Aircraft entered the hold at the GINLA waypoint as advised by the Controller.

The flight crew executed the required checklists, including the non-normal checklist for loss of hydraulic system A, and decided to return to OMDB. The flight crew requested vectoring for the approach and landed safely on runway 12L. After vacating the runway via taxiway N8, the Aircraft taxied to parking stand E23.

The Air Accident Investigation Sector determines that the cause of the Incident was the rotation of the damaged No. 1 tire inside the left main wheel well that consequently caused damage to some components and lines of hydraulic system A, and the subsequent hydraulic system A pressure loss.

The Air Accident Investigation Sector identifies the contributing factors to the Incident as follows:

  • The intermittent operation of the left alternate antiskid valve, which most probably allowed the two wheels of the left main landing gear to enter the wheel well while still spinning.
  • The rotating wheel peeled portion of the center tire tread rib of the number 1 wheel did not operate the frangible fitting, resulted in a continuation of the left gear retraction, allowing the number 1 tire peeled portion to damage the hydraulic system A and other parts located in the vicinity.

In this Final Report, the AAIS issued two safety recommendations addressed to the Operator, one to the General Civil Aviation Authority (GCAA), and one to the Aircraft manufacturer.

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1.1 History of Flight

On 8 August 2017, a flydubai Boeing B737-800, registration mark A6-FDS, operated a scheduled passenger flight FZ081, from Dubai International Airport (OMDB1) to Bahrain International Airport (OBBI2). There were a total of 35 persons on-board the Aircraft: 29 passengers, two flight crewmembers, and four cabin crewmembers. The Commander was the pilot flying (PF) and the Co-pilot was the pilot monitoring (PM). The Aircraft was pushed back from its parking stand E23 at about 0646 UTC. A long pushback was instructed since there was another aircraft being pushed back on the lefthand side of the Aircraft on parking stand E22. Taxi instructions were given to the flight crewmembers for a departure from runway 12R. The Aircraft was planned for a standard instrument departure runway 12R, via SITAT 2G. Before the departure, the flight crewmembers selected flaps 1. The take-off weight was approximately 54.6 tons, and the calculated rotation speed (VR) was 134 knots. The flight crew set the selected altitude to 4,000 feet (ft) as per OMDB standard instrument departure (SID). The Aircraft lined up on runway 12R for departure at about 0713. After take-off clearance was given, the flight crewmembers commenced the takeoff at about 0715. After the Aircraft lifted off the runway, the landing gear was selected up, and at about 250 ft radio altitude while climbing, the Aircraft experienced a loss of hydraulic system A, at 0716:35. The Aircraft continued climbing to the selected altitude. At 0716:55, Dubai air traffic Controller (ATC) Tower instructed the Aircraft to change to Dubai ATC Departure on frequency 121.025 MHz. At 0717:18, when the Aircraft was just passing 1,450 feet radio altitude, the Co-pilot informed Departure that the Aircraft had a hydraulic system problem, and requested to go to a hold area that was convenient for ATC. Departure Controller instructed the flight crew to continue the climb as per the SID. The Co-pilot then requested to maintain altitude at 4,000 feet, which was approved by the Departure Controller. The Departure Controller enquired whether the Aircraft could hold at GINLA, which was confirmed by the flight crew. Subsequently, at 0720:51, Departure Controller instructed the Aircraft to turn left and proceed directly to the GINLA waypoint, which was then performed by the flight crew. The Commander asked the Departure Controller for confirmation as to whether the Aircraft, at inbound GINLA, needed to turn right to enter the hold, which was confirmed by the Controller. The flight crew executed the applicable checklists, including the non-normal checklist for loss of hydraulic system A. At 0722:24, Departure asked the Aircraft an estimated time for holding at GINLA. The Commander responded that it will take a maximum of 10 minutes. At 0724:25, the Commander informed the cabin crew about the loss of the hydraulic system, and the extended landing gear. He also informed the cabin crew that the Aircraft would return to Dubai as a precaution in about 10 minutes. Thereafter, the Commander contacted the Operator’s network control center (NCC) and informed them that the Aircraft was at an altitude of 4,000 feet altitude about 30 nm from Dubai, had lost the hydraulic system A, and that the flight crew was planning to return to Dubai in 10 minutes. At 0733:16, the flight crew contacted Departure, and informed the Controller that the Aircraft was ready for the approach to runway 12L and requested radar vectoring. The Departure Controller then provided the requested radar vectoring.

At 0734:26, Departure informed the flight crew about the runway to be used, which was 12L and the QNH of 1,000 mbar. The Commander replied correctly, and informed that he expected to have some hydraulic fluid on the runway, which was acknowledged by Departure. At 0734:56, Departure vectored the Aircraft onto a heading of 360 degrees, and asked whether the flight crew were declaring an emergency. The Co-pilot acknowledged the instruction to take up a heading of 360 degrees, and he declared an emergency based on the decision of the Commander. The Departure Controller acknowledged the declaration of an emergency. At 0735:19, Departure instructed the Aircraft to contact ATC Approach on frequency 127.9 MHz. The Aircraft then contacted Approach and informed the Approach Controller that they were at 4,000 feet on a heading of 360 degrees. The Approach Controller instructed the flight crew to maintain the same altitude and heading. Subsequently, the Controller provided radar vectoring for the approach, and instructed the flight crew to report when established on the ILS for runway 12L. At 0737:35, the Aircraft turned right onto a heading of 030 as instructed by the Approach Controller. At 0738:24, Approach instructed the Aircraft to descend to altitude 2,000 feet, and subsequently the Aircraft started to descend. At 0738:55, Approach instructed the Aircraft to turn right to UKRIM waypoint for ILS runway 12L and to report when established, which was acknowledged correctly by the Co-pilot. At 0739:45, the Aircraft contacted Approach and informed that the Aircraft was being slowed and that the flight crew would extend the gear manually. The Approach Controller enquired about the required speed on final, which the flight crew advised would be 140 knots. The Commander contacted the NCC, and advised that the Aircraft would land within 5 minutes. He requested NCC to inform the Operator’s maintenance department about the fluid quantity of hydraulic system A, which was zero, and the 10 psi system pressure. Subsequently, the flight crew completed the manual gear extension checklist.

At 0744:41, Approach Controller asked the flight crew to confirm that the Aircraft was established on the glideslope. This was confirmed. Subsequently, Approach instructed the Aircraft to contact Dubai Tower on 118.75 MHz. The Co-pilot contacted the Tower, at 0745:11, and confirmed that the Aircraft was fully established on the ILS for runway 12L, with a distance to go of 6.2 nautical miles. The Tower provided wind information, cleared the Aircraft to land, and informed the flight crew to vacate the runway at any exit on the left as required. The Aircraft landed uneventfully on Runway 12L at 0748:03, vacated the runway via taxiway N8, and taxied to parking stand E23. The engines were shut down at 0757:47.

1.3 Damage to Aircraft

After the landing, the Aircraft was inspected, and it was found that the No. 1 tire (left outboard tire) tread center rib had completely peeled from the tire. However, the tire was still inflated (figure 4).

A6-FDS Damaged Tyre

The Aircraft sustained damage to the APU fuel drain line, landing gear and hydraulic parts (see Appendix 1) inside the left main wheel well, and minor damage to structural parts adjacent to that area. Scattered hydraulic fluid leakage in the left main wheel well was found after the Aircraft landed (figure 5).

2. Analysis

The Investigation collected data from various sources for the purpose of determining the causes and contributing factors that led to the Incident. This Analysis covers the tire damage and the consequences, the effect on Aircraft systems due to the loss of hydraulic system A and the related Operator’s procedures. This part of the Report explains the contribution of every investigation aspect to the Incident. The Analysis also contains safety concerns that may not be contributory to the Incident but are significant in adversely affecting safety.

2.1 The Damaged Tire

Although the No. 1 landing gear tire was damaged during takeoff, it was still inflated until the Aircraft returned to the parking stand after the flight. The tread pieces were not found after the Incident. It is most probable that the debris was lost from the wheel well, when the left main gear was extending, after it had not reached the up and locked position. For further examination, the damaged tire was sent to the tire manufacturer, who concluded that the tear-lines visible on the top belt ply, and the trace of cut/damage on the center inner tread groove opposite the serial side in sector 5 exhibited marks of a possible initiation site for the loss of the center tread rib. As the missing tread pieces were not returned, the tire manufacturer stated that they could not determine the root cause of the center tread rib loss. It is important to locate all tread parts or debris of the damaged tire in the case of tread separation, because these are the parts, which were in contact with the ground when the damage occurred. The trace of a cut with tear-lines on the top belt ply that extended out of the cut was possibly caused by a foreign object debris (FOD) or rough ground surface. However, the Investigation could not determine where, when and how the cut damage had occurred. There was no evidence that any FOD was present on the runway during the take-off, nor on other aircraft takeoff and landing, prior to or after the take-off of the Incident Aircraft. However, it is possible that the cut occurred on a previous flight and, was not detected on the inspections prior to the Incident. Premature signs of tire internal components separations, such as bulges, localized rubber splits, or uneven wear, may be a sign of delamination damage. In that case, it is important to remove tires from service when any evidence of separation is observed before a flight. During ground roll and high-speed rotation, even small areas of tread separation may grow into partial or full tread rubber loss. However, there was no evidence that the No. 1 tire had internal components separations prior to the flight. The tire had been re-treaded twice prior to the Incident. It had been installed on the Aircraft 13 days after the second re-treading that was carried out by the tire manufacturer according to the manufacturer’s specifications. The tire had been in service for 66 cycles prior to the Incident. Although the Investigation could not determine the cause of the tire damage, the Investigation believes that the Operator should include re-treaded tires in its reliability program, and to ensure that on-wing monitoring of nose and main wheel tire pressures is sufficiently robust to ensure TSO specifications for re-tread tire integrity are maintained throughout the tire operational spectrum. The reliability program should also consider the operational and environmental conditions. A non-destructive test (NDT) inspection (shearography) is required for tire re-treading to identify any abnormalities. According to the Boeing Service Letter 737-SL-32-128-A – Tire Retreading Recommendations, the re-treaded tires are recommended to go for NDT checks for an extent dependent on the re-tread level. Boeing recommended that the NDT inspection covers: the crown area for retread levels 0, 1 and 2; and the whole tire (bead-to-bead) for level 3 and beyond. Following the Incident, the Operator has taken a safety action by ensuring that the contracted re-treading agencies perform the required NDT checks dependent upon the retread levels as suggested by the Aircraft manufacturer and issued through Boeing Service Letter 737-SL-32-128-A, in addition to inspection requirements raised by the reliability data. The Investigation also reviewed the flight data prior to and during the takeoff. There was no evidence that the operations resulted in a significant lateral acceleration, which could have caused a substantial incremental side load on the tires, and could have contributed to the tire damage. Recent ramp inspections of other Operator’s aircraft conducted by the GCAA prior to the Incident, indicated that there were no issues related to tires observed during the inspections. The pressure of all tires was also checked after the Incident, and it indicated that two tires were within limits, and four tires were above the maximum limit. In order to improve detection of tire pressures, the Investigation recommends that the GCAA ensure that air operators on-wing monitoring of nose and main wheel tire pressures is sufficiently robust to ensure TSO specifications for re-tread tire integrity are maintained throughout the tire operational spectrum.

2.2 The Loss of Hydraulic System A and the Effect on Aircraft Systems

The hydraulic system of the Aircraft was designed with built-in redundancy. When the Aircraft lost hydraulic system A, system B and the Standby system automatically and manually provided the required hydraulic power to essential systems. The affected systems on the Aircraft when the hydraulic system A lost, were: autopilot A; flight spoilers (two on each wing); normal landing gear extension and retraction; ground spoilers; alternate brakes; No. 1 engine thrust reverser normal hydraulic pressure; and normal nose wheel steering. During the takeoff and until the Aircraft lost hydraulic system A, the autopilot was disengaged. The Commander selected autopilot A, which was engaged for approximately 5 seconds, when the Aircraft was climbing through 1,750 feet radio altitude. The Investigation believes that the five seconds engagement of autopilot A was due to the residual hydraulic pressure. The Commander realized that autopilot A was inoperative, and therefore requested the Co-pilot to select autopilot B. Some of the damaged components were related to the electric motor-driven pump, and to the engine driven pump of hydraulic system A, which caused the complete loss of the hydraulic fluid in the system and a total loss of system pressure. The Aircraft had eight flight spoilers, four on each wing. Four of the eight flight spoilers were inoperative as the result of the loss of hydraulic system A. Consequently, the effectiveness of Aircraft’s roll rate reduced, since the flight spoilers also supplemented roll control in response to control wheel inputs. The Investigation compared the roll rate of the Aircraft between the Incident flight and the previous flight and found that the maximum roll rate of the Aircraft, with autopilot engaged, decreased by about 50% due to the loss of hydraulic system A. The failure of the four of the eight spoilers might also have reduced the effectiveness of the speedbrake. The speedbrake was armed for the landing, and since hydraulic system B was available, the four system B flight spoilers deployed. The speedbrake lever automatically moved to the UP position. During the landing, when the Aircraft touched down, the Aircraft’s auto brake system functioned normally since the normal brake system, powered by the hydraulic system B, was operative. The Aircraft lost the alternate brake system due to loss of hydraulic system A. The loss of hydraulic system A during landing gear retraction caused the left main gear and the nose gear to remain in transition and they did not reach the up and locked position. The right main gear reached the up and locked position. Since the landing gear lever was at the up position, the landing gear warning of the left and nose gear appeared. The left and nose gear thereafter fell to the down and locked position. The nose gear most probably did not reach the up and locked position because the loss of hydraulic system A occurred before the nose gear reached its up and locked position. The flight crewmembers performed the manual landing gear extension since the normal extension procedure could not be applied due to the hydraulic system A loss and the failure to release the up-locks. The manual gear extension handles were activated by the Co-pilot, which released the up-locks and allowed the right main landing gear to fall into its locked position. During the landing, the four ground spoilers were inoperative since they were powered by hydraulic system A, as recorded on the flight data. It was also recorded that the spoilers 3 and 10 were deployed on the touchdown, which also means that flight spoilers 5 and 8 (not recorded) were also deployed. Other flight spoilers that were powered by hydraulic system A (2, 4, 9 and 10) were inoperative. This condition resulted in a longer landing distance. The No. 1 engine thrust reverser was inoperative due to the loss of hydraulic system A. Although the Standby hydraulic system was available to power the No.1 engine thrust reverser, the Commander chose not to apply the thrust reversers for the landing since he was aware that by using the Standby system, the slow rate of the deployment and retraction of the affected thrust reverser could result in thrust asymmetry during landing. The Investigation compared the landing distance between the Incident flight and one of the previous flights, with the same Aircraft landing configuration (with flaps 30) and similar flight condition (in groundspeed, and almost the same wind direction and speed). The landing distance was measured from when the Aircraft touched the ground, all spoilers deployed in the up position for the other flight of the Aircraft and only operative spoilers deployed for the Incident flight, until the point that the deployed operative spoilers retracted. Auto brakes were applied on both flights. Adjustment of the distance was made since there was a difference in the weight of approximately 5.4 ton, as required by the airplane flight manual (AFM). This revealed that the landing distance of the Incident flight was increased by approximately 48 percent due to the loss of hydraulic system A and the decision not to use the thrust reversers on both engines. Normal nose wheel steering was inoperative due to the loss of hydraulic system A. Consequently, the Commander opened the guarded switch and selected the nose wheel steering switch to ALT, which provided nose wheel steering powered by hydraulic system B. After the landing, no issues were found on the application of the nose wheel steering during taxiing the Aircraft to the parking stand. All the evidence and the analysis shows that the redundancy of the Aircraft’s hydraulic system and the affected systems after the loss of hydraulic system A were operating as designed.

2.3 Flight Operations

Both flight crewmembers operated the flight in accordance with the Operator’s standard operating procedures (SOP). They carried out all necessary checklists, including the non-normal checklist for loss of hydraulic system A. After the loss of hydraulic system A, the Commander requested to enter the hold. His intention was to have sufficient time for both flight crewmembers to understand the problem, make decisions, conduct the necessary failure briefing, and take appropriate action. The Commander also briefed the cabin crew on the situation and the decision to return to OMDB. The GINLA waypoint provided by the ATC was not a dedicated holding pattern. The Investigation believes that the provided holding area to the Aircraft including the vectoring did not create significant workload issue(s) for the ATCs to manage departures and arrivals at OMDB, since there was no evidence of any ATC issues at the period of the Incident flight. After the loss of hydraulic system A, the flight crewmembers had no indication that the Aircraft had experienced tire damage. Therefore, the flight crewmembers did not inform ATC about the possibility of debris on the runway. The external and internal cockpit communications were conducted in accordance with the Operator’s procedures. The Commander was the pilot flying, however, more than once, he transferred the control of the Aircraft to the Co-pilot, when he performed communications with the ATC, the Operator’s network control center (NCC), and the cabin crew. The Commander performed these as considered necessary. Nevertheless, high alertness and situational awareness were maintained by both flight crewmembers throughout the flight.

2.4 Brake and Protection on Landing Gear Retraction

The operational test of the gear retract braking performed by the Operator after the Incident, it revealed that the left main landing gear brake system functioned normally, which means that the hydraulic pressure had been applied. The landing gear retraction of the incident flight started about six second after the Aircraft lifted off. A portion of the center tread rib of no.1 tire probably peeled off during the takeoff roll and before the landing gear retraction. After the landing gear lever was set to the up position, the brake pressure increased from 30 to 490 psi, indicating that brake pressure from the metering valve was normal for landing gear retraction. Approximately seven seconds later, a low-pressure indication of hydraulic system A appeared, and the brake pressure decreased to the original pressure value, 30 psi, and after several seconds reduced to zero. During the seven second transition of the landing gear retraction, the alternate brake selector valve sent landing gear retract pressure (up pressure) to the alternate brake system, which operated the alternate brake metering valves and sent pressure to the brakes. When hydraulic system A was lost, the right main landing gear had fully retracted into the wheel well and both wheels had stopped rotating. The portion of the peeled center rib of No.1 tire tread did not remove the lever from the frangible fitting since it was found attached to the frangible fitting after the Incident. Therefore, the frangible fitting did not open/operate, and the retraction of the left main landing gear continued. When the left wheel went into the wheel well, the rotating peeled portion of the center tread struck and damaged the components and lines related to hydraulic system A, and other components in the left wheel well. Consequently, the Aircraft lost its hydraulic system A. The loss of hydraulic system A caused the left main landing gear actuator to stop extending, which resulted in the left main landing gear not reaching its up and locked position. Almost at the same time, the right main landing gear reached its up and locked position. The nose gear retraction stopped and did not reach its up and locked position when hydraulic system A was lost. Since the frangible fitting did not open or operate, no hydraulic flow went out through the frangible fitting, consequently, the volume fuse did not close the gear up pressure line and the flow limiter never detected the flow limit, since it never reached the limit. This condition resulted in an inability to stop the loss of hydraulic system A fluid. The other outcome was that the gear retraction transition continued, and the gear entered the wheel well. Subsequently, the retraction stopped when the hydraulic system A pressure was lost. Had the portion of the peeled center rib of the No.1 tire tread removed the lever of the frangible fitting, the loss of hydraulic system A might have been prevented and the hydraulic pressure might have been available for other Aircraft systems that were powered by hydraulic system A. Furthermore, the retraction of the left landing gear may have stopped before the gear entered the wheel well. The duration of a main gear wheel to stop spinning after the liftoff depends upon some variables, such as: the aircraft speed when lifting off; the time interval between the aircraft lift off and gear retraction; the aerodynamic forces on the tire; the friction of the wheel bearing; the availability of hydraulic pressure from the landing gear retract pressure to the alternate brake metering valves; and the metered hydraulic pressure from the alternate brake metering valve to the brakes. Based on the flight data available from previous flights of the Aircraft the retraction time was approximately seven seconds, from the gear up lever setting to the up and locked position of the main landing gear, which was also the case for the right main landing gear in this Incident. For a normally operating brake control system, the manufacturer designed the gear retract braking function to have completely stopped the main gear wheel/tire assemblies from spinning within approximately 1 second after the landing gear lever is moved to the UP position. This includes the time for the landing gear selector valve to provide pressure to the alternate brake metering valve, brake assemblies, and alternate antiskid valve, and the time for the wheel/tire assembly rotation to be stopped. This ensures that the main gear wheel/tire assemblies are not spinning when they begin to enter the wheel wells. The antiskid function is inhibited for 12.5 seconds after the landing gear lever is moved to UP to prevent an antiskid release of the gear retract brake pressure anytime during the gear retraction phase. However, in this event, the Investigation believes that the alternate antiskid valve may have been operating intermittently or in a degraded mode, resulting in an inconsistent and/or incomplete gear retract braking function for the left main gear wheel/tire assembly. The frangible fitting is essential in preventing damage to the components in the wheel well, as designed in accordance with the certification specification of FAR 25.1309(b). The Investigation could not determine why the frangible fitting did not open or operate, since the lever was still attached following the Incident. Therefore, it is recommended that a risk assessment exercise be carried out by the Aircraft manufacturer on the protection mechanism system of components in the wheel wells considering the functionality of the frangible fitting in a situation of tire damage during landing gear retraction. On 25 July 2017, two weeks before the Incident, the left gear retraction brake system was found unserviceable. The braking did not stop the rotation of the left main gear wheels during retraction. However, the Aircraft was released for service in accordance with the applicable minimum equipment list (MEL) and an acceptable deferred defect (ADD) was raised. On the same day, another acceptable deferred defect was raised regarding the required cleaning on the left main wheel well area due to dirty black rubber traces in the left main gear wheel well. One week later, on 1 August 2017, the related acceptable deferred defect of the unserviceable retraction brake was closed, after the wheel well brake-metering valve was replaced. However, the acceptable deferred defect of the required cleaning on the left main wheel well area was still open until the Incident. When an ADD of an unserviceable brake system exists, the Operator should ensure the suitability or fitness of the tire by performing additional checks on the tire. A clean mainwheel well could provide an indication of the function of the gear retraction brake system, in the case of damaged tire(s), as occurred before the Incident on 25 July 2017. Therefore, the Investigation recommends establishment of appropriate mechanisms to ensure that technical washing of the nose and main wheel well compartments is undertaken with sufficient frequency to allow the integrity of aircraft systems to be assessed sufficiently and continuously. In this regards, a policy is to be established not to defer any maintenance corrective actions without proper documented reference.

3. Conclusions

3.1 General

From the evidence available, the following findings, causes, and contributing factors were made with respect to this Incident. These shall not be read as apportioning blame or liability to any particular organization or individual.

To serve the objective of this Investigation, the following sections are included in the Conclusions heading: -

  • Findings- are statements of all significant conditions, events or circumstances in this Incident. The findings are significant steps in this Incident sequence but they are not always causal or indicate deficiencies. -
  • Causes- are actions, omissions, events, conditions, or a combination thereof, which led to this Incident. -
  • Contributing factors- are actions, omissions, events, conditions, or a combination thereof, which, if eliminated, avoided or absent, would have reduced the probability of the Incident occurring, or mitigated the severity of the consequences of the Incident. The identification of contributing factors does not imply the assignment of fault or the determination of administrative, civil or criminal liability.

3.2 Findings

3.2.1 Findings relevant to the Aircraft

(a) The Aircraft was certificated, equipped, and maintained in accordance with the requirements of the Civil Aviation Regulations of the United Arab Emirates.

(b) The Aircraft was airworthy when dispatched for the flight.

(c) The Operator reported that several components were removed and tested following the event. The left antiskid valve was found faulty and may have been operating intermittently.

(d) The Aircraft had one acceptable deferred defect (ADD) related to cleaning the left main wheel well area before the flight. The ADD was still open on the Incident flight.

(e) The No. 1 tire experienced a partial center rib peel during take-off.

(f) During the landing gear retraction, the portion of the peeled center rib of No.1 tire tread did not remove the frangible fitting, which resulted in a continuation of the left main gear retraction.

(g) When the No.1 left wheel went into the wheel well, the rotating peeled portion of the center tread damaged components and lines related to hydraulic system A, consequently, the Aircraft lost its hydraulic system A.

(h) The nose and left main landing gear stopped retracting before reaching the up and locked position, while the right main landing gear reached its up and locked position.

(i) It is most probable that the left alternate antiskid valve was intermittently operating correctly and allowed the spinning wheels of the left main landing gear to enter the wheel well.

(j) The Aircraft returned to the departure airport and landed uneventfully.

(k) There was no evidence of foreign object debris before takeoff.

(l) The tread pieces or debris were not found following the Incident.

(m) The damaged No. 1 tire remained inflated.

(n) The tire manufacturer could not determine the root cause for the center tread rib loss.

3.2.2 Findings relevant to the flight crewmembers

(a) The flight crewmembers were licensed and qualified for the flight in accordance with the existing requirements of the Civil Aviation Regulations of the United Arab Emirates.

(b) Both flight crewmembers operated the flight in accordance with the Operator’s standard operating procedures (SOP). They carried out all necessary checklists, including the non-normal checklist, for the loss of hydraulic system A.

(c) Both flight crewmembers maintained sufficient situational awareness during the Incident flight.

3.2.3 Findings relevant to flight operations:

(a) The Commander requested holding in order to obtain sufficient time for both flight crewmembers to understand the problem, make decisions, conduct the necessary failure briefing, and take the necessary action.

(b) The Commander informed the cabin crew manager, and also the Operator’s network control center (NCC), of the hydraulic system A failure and the plan to return to the departure airport.

3.2.4 Findings relevant to air traffic control:

(a) The provided holding area to the Aircraft, including the vectoring to landing, did not create significant workload issue(s) for the ATCs to manage departures and arrivals traffic at the departure airport.

3.2.5 Findings relevant to the Operator

(a) There was no time limit applicable to correcting the open deferred defect pertinent to cleaning the main wheel well area.

3.3 Causes

The Air Accident Investigation Sector determines that the cause of the Incident was the rotation of the damaged No. 1 tire inside the left main wheel well that consequently caused damage to some components and lines of hydraulic system A, and the subsequent hydraulic system A pressure loss.

3.4 Contributing Factors to the Incident

The Air Accident Investigation Sector identifies the contributing factors to the Incident as follows:

(a) The intermittent operation of the left alternate antiskid valve, which most probably allowed the two wheels of the left main landing gear to enter the wheel well while still spinning.

(b) The rotating wheel peeled portion of the center tire tread rib of the number 1 wheel did not operate the frangible fitting, resulted in a continuation of the left gear retraction, allowing the number 1 tire peeled portion to damage the hydraulic system A and other parts located in the vicinity.

4. Safety Recommendations

4.1 General

The safety recommendations listed in this Report are proposed according to paragraph 6.8 of Annex 13 to the Convention on International Civil Aviation, and are based on the conclusions listed in part 3 of this Report; the Air Accident Investigation sector expects that all safety issues identified by the Investigation are addressed by the concerned organizations.

4.2 Safety Action

Following the Incident, the Operator has taken a safety action by ensuring that the contracted tire retreading agencies implement the tire re-tread reliability-based inspection (including shearography) requirements and the acceptance criteria, as suggested by the Aircraft manufacturer and issued through Boeing Safety Letter 737-SL-32-128-A.

4.3 Final Report Safety Recommendations

The Air Accident Investigation Sector recommends that:

4.3.1 flydubai SR09/2018 To include re-treaded tires in its reliability program, and to ensure that on-wing monitoring of nose and main wheel tire pressures is sufficiently robust to ensure TSO specifications for re-tread tire integrity are maintained throughout the tire operational spectrum. SR10/2018 To establish appropriate mechanisms to ensure that technical washing of the nose and main wheel well compartments is undertaken with sufficient frequency to allow the integrity of aircraft systems to be assessed sufficiently and continuously. In this regards, a policy is to be established not to defer any maintenance corrective actions without proper documented reference.

4.3.2 The General Civil Aviation Authority of the United Arab Emirates SR11/2018 To ensure that air operators on-wing monitoring of nose and main wheel tire pressures is sufficiently robust to ensure TSO specifications for re-tread tire integrity are maintained throughout the tire operational spectrum.

4.3.3 Boeing SR12/2018 To carry out a risk assessment exercise on the protection mechanism system of components in the wheel wells considering the functionality of the frangible fitting in a situation of tire damage during gear retraction.

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