After departing Brisbane en-route to Singapore, the crew of the Boeing 777-212ER aircraft heard and felt two thumps through the airframe and noticed a severe vibration indication of the right engine. The vibration subsided before re-occurring moments later with an increase in the engine's exhaust gas temperature also indicated. The crew conducted an in-flight engine shutdown and requested Air Traffic Control clearance to divert to Darwin where an uneventful single-engine landing was carried out.
An inspection conducted by ground engineers at Darwin found a stage-1, variable stator vane (VSV) control lever broken on the right engine. A boroscope inspection of the engine interior was then carried out with a number of compressor blades found damaged. The engine was removed from the aircraft and sent to an overhaul facility for disassembly and evaluation.
The Rolls-Royce Trent 800, was a triple spool turbofan engine. Its construction consisted of a single-stage low pressure fan connected to a five-stage low pressure turbine. An eight-stage intermediate pressure compressor (IPC) connected to a single stage intermediate pressure turbine and a six-stage high pressure compressor (HPC) connected to a single-stage high pressure turbine (see fig 1).
To maintain maximum efficiency during all power settings, the airflow through the engine needed to be controlled to prevent stalling or surging. This was achieved by a single stage of variable inlet guide vanes installed between the fan and the first stage of the IPC and two stages of variable stator vanes, VSV stage 1 and VSV stage 2 (VSV-1 and VSV-2) installed between IPC stages, one/two and two/three respectively. Each Variable Stator Vane was attached to a lever that transferred the linear input from the controlling actuators and unison rings to a rotational movement of the vane (see fig 2). These levers consisted of an arm and connecting pin (see fig 3).
During engine start, these vanes would have been in their most closed position with internal engine bleed valves open. As the power was increased, the bleed valves would close and the vanes move toward their full open position allowing optimum airflow through the engine.
Post incident engine inspection
Prior to disassembly, the engine's exterior was inspected with the broken VSV-1 lever identified as being in the number-28 position. No evidence was observed of bird impact or other external defects. Rigging and clearance checks of the VSV control system were carried out with no anomalies found. The remaining VSV-1 and VSV-2 levers were then removed and crack tested using a dye penetrant inspection. No evidence of cracking was found on any of those levers.
After separating the engine-to-modular level, the IPC and HPC modules were completely disassembled for a detailed inspection.
Removal of the IPC casing revealed six stage-2 blades displaying soft body impact damage 1 resulting in bending of the blades. On two of those blades the corners had also detached. Three other blades displayed hard body impact damage 2 with cuts and nicks (small cuts) on their surfaces. Two blades with minor nicks were found in stage 3, with only one blade in stage 5 showing nick damage. All of the stage-8 blades displayed hard body impact damage on their trailing edges, a few blades also having nicks on their leading edges. There was no evidence of damage to the disc material adjacent to the blade roots on any of the eight stages.
The IPC case lining was examined with only minor damage evident. The number-28 VSV 1 vane, had a wear mark on the leading edge lower corner with a noticeable worn stepped area on its horizontal surface above its base. The adjacent number-29 vane had a contact mark at a point mid span on the vane and one on its base. The remaining VSV-1 and all the VSV-2 vanes were found to be undamaged. Dark deposits were evident around the base of each VSV. These deposits formed a black ringed area around all except for the number-28 vane where the mark was crescent shaped.
When the number-28 and 29 vanes were positioned so that the wear marks on both vanes aligned, the number-29 vane was found to be in its normal full open position while the number-28 vane was noted as sitting beyond its normal closed position. The dark crescent area around the base of the number-28 vane also coincided with the angle of the vane's root. When the number-29 vane was moved to the closed position it was seen to nudge the number-28 vane up towards its normal closed position.
Failed VSV-1 lever
The Australian Transport Safety Bureau (ATSB), conducted a metallurgical examination of the failed number-28, VSV-1 lever (see fig 3). The examination found that:
'The lever had fractured transversely through the end of the arm section, at a location coincident with the riveted connection to the actuator pin. The fracture path followed a uniform arc, extending from one side of the arm to the opposite and intersecting the pin connection at the centre. A prominent track mark had developed on the underside of the arm where the relative movement between the separated arm and the pin flange had produced appreciable wear. On close inspection, the fracture path appeared to intersect the bore of the rivet hole with a slight upward 'dishing' of the arm section beneath the rivet head'.
The examination also determined that:
'During riveting, the expansion of the rivet shaft could induce tensile stresses within the bore of the rivet hole if the diameter was insufficient to allow for the expansion. Tensile stresses of this nature would be expected to predispose the lever arm to the initiation and propagation of fatigue cracking'.
The examination of a further four VSV-1 levers was conducted, with welding and partial fusion between the lever and connecting pin evident, and varying degrees of cracking also evident on all four levers. For the full technical report refer to attachment A.
A further investigation by the engine manufacturer, identified the presence of a double-sided chamfer to the lever holes on a small number of levers. This removal of material during the lever manufacture may have led to the overheating and partial welding of the lever material during the rivet forming.
In the HPC, all of the stage-1 blades displayed severe hard body impact damage with one blade found to have failed, detaching above the blade root. Stages 2 to 6 showed hard body impact damage to varying degrees on all the blades.
Close examination of the failed HPC stage-1 blade, found a chipped area in the leading edge, with the fracture surface revealing a number of crack progression marks indicating that the failure was progressive over a number of cycles and not instantaneous. The exact number of cycles required to fail the blade could not be determined (see fig 4).
The engine commenced service in December 1998 and had completed a total of 8923 hours and 2373 cycles at the time of this incident. Its service history showed that on 8 April 2001, a routine boroscope inspection detected damage to a number of IPC stage 2 blades in the form of bending and curling to their tips. This damage was assessed to be within the manufacturer's allowable limits so the engine remained in service. On 17 October 2001, a substantial shift in the turbine gas temperature (TGT) was detected giving a warning that the engine's efficiency had deteriorated significantly. A boroscope inspection was carried out on the engine with only the previously recorded IPC stage-2 bent blade damage found. No other damage was observed on the engine. Checks of the air system, engine bleed air and monitoring systems were carried out, however the reason for the TGT shift could not be determined.
Previous VSV lever failures
Although the manufacturer had not experienced previous failures of VSV-1 levers, failures of VSV-2 levers had been recorded on two separate occasion. On those occasions the connecting pin's had fretted through the body of the lever due to inadequate riveting during the manufacturing process. The result of the levers failing was the closure of their associated variable stator vanes, which created a disruption to the airflow behind them. The vibration subsequently experienced by the blades passing the area resulted in fragments of disc material breaking off and migrating through the engine, damaging blades further down stream. These failures were indicated by a progressive increase in the engine's TGT over periods of 2 to 4 weeks.
The manufacturer issued a service bulletin, RB211-72-D516 to all operators recommending an inspection of, 'the six VSV-1 and VSV-2 levers either side of the actuating mechanism control rod connection, for significant relative movement between the lever and connecting pin'.
The soft body damage identified as bending on the IPC stage-2 blades by the boroscope inspection on 8 April 2001, was consistent with that incurred by the engine ingesting a soft bodied object such as a bird or ice. These defects were determined to be within serviceable limits, and as such would not have failed unless additional abnormal forces were applied to them.
There was no evidence of any system faults or additional internal deterioration of the engine after the increase in TGT was detected on 15 October 2001. It is possible that the increase was the result of the number-28 vane moving to the closed position after the lever failed. The nudging of the number-28 vane towards its normal closed position by the number-29 vane during engine shut down may have been enough to allow the failed lever to assume its normal position, thereby escaping easy detection.
The ATSB's technical analysis report determined that the failure of the lever was probably a product of a progressive fatigue cracking mechanism. The observed fracture features suggested that the crack initiation occurred from the connecting-pin, hole.
The bias in the wear mark to one side of the failed lever combined with the contact points evident on the number-28 and 29 vanes indicated that during engine operation, the number-28 vane had remained in a predominantly closed position.
As found with previous VSV lever failures, the out of sequence position of the vane created turbulence in the airflow. This would have been felt on the IPC stage-2 blades inducing abnormal loads. Two of the bent IPC stage-2 blades were unable to tolerate that excitation and as a result their blade tips failed. The released sections of blade then ricocheted within that stage before being projected through the engine, impacting blades in other stages down stream.
As the sections of blade and accumulated debris passed through the engine, a piece of material impacted the leading edge of a HPC stage-1 blade, chipping and cracking it. The crack then progressed to a point where the blade failed and detached.
Although it was possible for the IPC stage-2 blade tips to fail as a result of bird or ice ingestion during the incident flight, the lack of supporting evidence, and the failure mechanism of the HPC blade did not support such an event.
- A number of IPC stage-2 blades were found to be damaged during a boroscope inspection on 8 April 2001 but remained in service on the engine.
- Fatigue cracking of the VSV-1 lever led to its failure resulting in the closing of the number-28 Variable Stator Vane.
- The closure of the number-28 Variable Stator Vane created a disrupted airflow, which acted on the passing blades.
Local safety action
As a result of this incident the operator has proceeded to replace all VSV levers of the same manufacture as the failed item.
The engine manufacturer has amended the 'IPC tip bend acceptance criteria' text in the aircraft maintenance manual and issued a revision to service bulletin RB211-72-D516, extending the range of the inspection to include all VSV-1 and VSV-2 levers. The engine manufacturer has also designed a strengthened VSV lever that has been certified for use under service bulletin RB211-72-E042.
As a result of this occurrence the Australian Transport Safety Bureau issues the following safety recommendations:
1. Recommendation R20030002
The Australian Transport Safety Bureau recommends that Rolls-Royce plc revise service bulletin RB211-72-D516 to highlight the potential for cracking failure between the lever and connecting pin of the Variable Stator Vane lever assemblies, and ensure that inspections contained within this service bulletin adequately address this mode of failure.
2. Recommendation R20030003
The Australian Transport Safety Bureau recommends that the United Kingdom Civil Aviation Authority review Rolls Royce plc, Trent 800 engine inspection procedures for the variable stator vane lever assemblies and service bulletin RB211-72-D516, to ensure that they adequately address and manage the potential for cracking failure of the lever assemblies.
Technical Analysis Report
Boeing 777-212ER, 9V-SRE
1. FACTUAL INFORMATION
During a flight from Brisbane to Singapore, the crew of the Boeing 777-212ER aircraft noticed the onset of abnormal vibration levels and several 'thumps' from the right engine. While continuing to monitor the engine, the vibration and thumps recurred and the engine was subsequently shut down after the oil and exhaust gas temperatures rapidly increased. Following a diversion to Darwin, the crew conducted an uneventful single engine landing.
On initial inspection, maintenance personnel found a single fractured first-stage variable stator vane (VSV) control lever (figure 1). Later internal boroscopic inspection of the engine found significant levels of mechanical damage within the intermediate and high-pressure compressor stages.
The fractured lever and a selection of other levers from the first-stage VSV assembly were removed from the engine for examination by the ATSB.
Fig. 1 External view of the stage-one variable stator vane actuator ring and the single fractured lever (arrowed).
1.2 Visual examination and fractography (failed lever)
Initial inspection in the ATSB laboratory found the VSV lever had fractured transversely through the end of the arm section (figure 2), at a location coincident with the riveted connection to the actuator pin. The fracture path followed a uniform arc, extending from one side of the arm to the opposite and intersecting the pin connection at the centre (figure 3).
Fig. 2 Underside of the VSV lever removed from the engine.
Fig. 3 Underside of the VSV lever showing the curved fracture path and the wear mark produced by in-service movement after failure.
A prominent track mark had developed on the underside of the arm where the relative movement between the separated arm and the pin flange had produced appreciable wear. The effects of wear extended to the fracture surfaces themselves, which were heavily eroded and all fracture surface detail obliterated (figures 4 & 5). Apart from the fracture, the arm had sustained little other mechanical damage and showed no evidence of deformation or distortion associated with the failure.
Fig. 4 Arm section fracture surface showing degree of wear and loss of detail.
Fig. 5 Pin section fracture - adjoins the surface shown in figure 4.
On close inspection, the fracture path appeared to intersect the bore of the rivet hole, with slight upward 'dishing' of the arm section beneath the rivet head (figure 6). A clearance or gap was not evident between the pin shaft and the bore of the arm hole through which the pin shaft was riveted.
Fig. 6 Gap between the pin flange and the control arm produced by upward 'dishing' of the arm beneath the rivet head.
The examination did not show any evidence of binding or excessive friction between the actuator pin and the mating bushing, nor did any other component show significant indications of miss-installation or anomalous operation.
|Date:||18 November 2001||Investigation status:||Completed|
|Time:||1745 hours CST|
|State:||Northern Territory||Occurrence type:||Abnormal engine indications|
|Release date:||18 March 2003||Occurrence category:||Incident|
|Report status:||Final||Highest injury level:||None|
|Aircraft manufacturer||The Boeing Company|
|Type of operation||Air Transport High Capacity|
|Damage to aircraft||Nil|
|Departure point||Brisbane, Qld|