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During the take-off roll, as the Saab 340 aircraft reached about 100 kts, the crew heard a loud bang that was followed by loss of power from the right engine. The crew rejected the take-off and shut down the engine when the inter-stage turbine temperature increased to about 1,190 degrees Celsius. The air traffic controller confirmed the absence of fire or smoke and the crew returned the aircraft to the gate where the passengers disembarked.

An external examination by the aircraft operator's engineers found no evidence of damage to the engine or the intake from uncontained failure, case rupture, fire, or foreign objects. Two stage-4 air tubes were found broken and the exhaust centre body was missing. An internal boroscopic examination revealed that about one half of a single first-stage compressor blisk (bladed disk) blade had separated, and that the compressor and turbine blades sustained varying degrees of damage on the leading and trailing edges.

Due to the lack of authorised facilities in Australia, the operator sent the engine to its manufacturer in the USA for detailed examination and repair. The ATSB requested the National Transportation Safety Board (NTSB) of the USA to supervise the examination and provide a report to the Bureau. The ATSB also requested that the first-stage compressor blisk be returned for examination to determine the mode of blade failure.

The NTSB report confirmed that about one half of a single first-stage compressor blisk blade was missing and that all downstream engine components sustained varying degrees of damage consistent with the separated portion of the blade passing through the engine. The report also indicated that the engine was found to have been correctly assembled and that no deficiency was found that would have contributed to the separation of a single first-stage compressor blisk blade.

The only anomaly reported was a small difference in setting of the variable stator vane (VSV) assembly opening and closing angles, when compared to the build-up overhaul requirements. The manufacturer advised that such a difference could be expected, as it was reflecting normal in-service wear. The difference in the angles was not expected to affect the engine operation, but could reduce its stall margin.

The NTSB report further detailed damage to the power turbine shaft that was found intact, but containing an amount of black grainy material deposits at some galleries and seals. Commencing about 400 mm from the front end, the shaft also contained about a 200 mm wide strip of circumferential rub marks.

Engine and blisk history

The CT7-9B turbopropeller engine, serial number 785492, had accumulated 13,930.9 hours and 16,226 cycles since new and 4,428.2 hours and 4,718 cycles since overhaul. It was last overhauled in the UK in November 1998.

The engine design employed the "modular" concept with the engine consisting of the accessory, the core and the power turbine modules. The blisk, part number 6055T83G14, serial number GLHG4996, was re-installed into the engine core section module during the engine overhaul in November 1998 and remained a part of the engine since that time.

Since April 1999, the engine was installed into a number of aircraft operated by the same operator and was maintained in accordance with the requirements current at that time. In July 1999, the engine was shut down due to erratic operation and torque fluctuation. The subsequent examination found no fault, with the problem attributed to pilot training. In January 2001, the engine was removed due to noise coming from the gearbox. Records indicated that the hydromechanical unit was replaced and the engine was returned to service.

Blisk examination

Detailed examination of the failed first-stage compressor blisk is covered in Technical Analysis Report 39/01. [See associated tab above]

The examination revealed that one blade had separated about 29 mm from the root and that the remaining blades had sustained multiple impacts and distortions along the leading edges and tips.

The blade separation was found to have occurred as a result of fatigue cracking that initiated from corresponding transverse mid-span locations on both sides of the blade. The initial development and growth of the crack had been slow, extending over numerous hours and flight cycles. At approximately nine cycles before failure, the crack began to advance much more rapidly. The crack grew to a critical size and final overload fracture of the remaining section allowed separation of the outer blade section.

Many shallow cracks, not detected by non-destructive examination, were found adjacent to the primary fractures and in an adjacent blade that had not failed. Many of those cracks contained oxide and corrosion products attesting to their slow growth, and therefore, existence prior to the blade failure event. That presence tended to suggest that the failed blade was subjected to a reverse bending in the second or higher bending modes. No anomalies were found within the blisk material and manufacture.

Previous failures of the first-stage compressor blisk blade

The manufacturer reported that 12 similar events occurred during the engine's 17-year history. Investigation of the previous events led to the conclusion that the engine could experience a non-synchronous vibration that excited the power turbine shaft's first bending mode in the 10,300 to 14,500 RPM range. The power turbine shaft had a rated speed of 22,000 RPM.

The manufacturer reported that the onset of non-synchronous vibration could cause the bending power turbine shaft to contact and rub against the compressor tie rod, resulting in severe metal-to-metal rub between the components. It could also set up a vibratory stimulus, exciting reverse bending in the second and higher bending modes of the first-stage compressor blisk blades, resulting in high stresses at the blades' mid-span location and subsequent crack initiation and propagation. The relative positions of the individual components on the engine is shown at Attachment "A".

The onset of non-synchronous vibration was attributed to friction in the rotor system at the power turbine shaft forward spline due to either lack of lubrication, spline wear, misalignment or reduced damping at the number-2 bearing in the input drive assembly.

Following research into the non-synchronous vibration in the early 1990's, the manufacturer issued a number of service bulletins aimed towards alleviating the problem. The ATSB was advised that the subject engine's power turbine shaft, either during manufacture or at maintenance, had all requirements of the relevant service bulletins incorporated and that the investigated mid-span blade separation was the first such occurrence since June 1995.

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The manufacturer advised that previous research into the mid-span blade failures indicated that the failures were always associated with the presence of vibratory loads. The loads were found to have resulted from the higher levels of non-synchronous engine vibration that were evidenced by the power turbine shaft contacting and rubbing against the compressor tie rod.

The failed engine power turbine shaft was reported to be exhibiting evidence of rubbing against the compressor tie rod. The manufacturer advised that, although the rubbing was not as severe as observed during the previous first-stage compressor blisk mid-span blade separations, the required frequency, severity, and duration of the rub resulting in the blade separation was not known.

Previous research into the problem attributed onset of non-synchronous vibration to friction in the rotor system at the power turbine shaft forward spline due to either lack of lubrication, spline wear, misalignment, or reduced damping at the number-2 bearing in the input drive assembly. The NTSB engine examination report mentioned no deficiency in any of those areas.

On completion of the examination by the ATSB, the blisk was returned to the manufacturer. The ATSB requested to be informed of any further development in the investigation and research into the onset of the engine non-synchronous vibration by the manufacturer's specialists. Any results from that research will be published on the ATSB website at: www.atsb.gov.au.

 
  1. Onset of the engine non-synchronous vibration excited reverse bending of the first-stage compressor blisk blade in the second or higher bending modes.
  2. Presence of fatigue cracking in the mid-span location on a single first-stage compressor blisk blade led to blade separation.
  3. Damage to the downstream engine components by the separated portion of the blade resulted in a loss of engine power and increased inter-stage turbine temperature.
 
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Examination of a Failed Compressor Blisk - Saab Aircraft AB, SF-340B, VH-EKX

1. FACTUAL INFORMATION
1.1 Introduction

On 23 May 2001, the take-off of a Saab 340 aircraft (VH-EKX) was discontinued as a result of the right engine failing. To investigate the event, the Australian Transport Safety Bureau carried out an examination of the first stage compressor blisk (bladed disk) from the affected engine. An earlier disassembly of the engine had found the component damaged by the partial loss of a single aerofoil (blade) section. The examination attributed the downstream damage within the engine to the effects of the released aerofoil section.

1.2 Component history

The blisk component was installed within a General Electric CT7-9B turboprop engine; serial number 785492. At the time of failure, the engine had operated for a total of 13,931 hours and 16,226 flight cycles. The engine manufacturer indicated that the failed stage-one blisk (p/no. 6055T83G14, s/no. GLHG4996) had been fitted to the engine since new. A supplied specification indicated that the blisk was produced as a forging from a proprietary precipitation hardening stainless steel alloy, similar to UNS S35500 (AM355).

1.3 Visual examination and fractography

1.3.1 General condition
The leading edges and tips of the blisk aerofoil sections had been damaged by multiple impacts (figures 1 & 2). The single fractured aerofoil had separated transversely at a mid-span location, around 29 millimetres above the root transition (figure 3). The rear edges of the blisk hub (adjacent to the stage two compressor wheel coupling) showed appreciable erosion around the full circumference (figure 4). Heat tinting and other evidence indicating rubbing contact against the housing was present over approximately two-thirds of the blisk circumference (figure 5). The forward (convex) surfaces of the aerofoils were coated with an oily, black deposit which increased in density toward the blade tips (figure 6). Beneath the deposit, the surfaces exhibited a bright, lustrous finish, typical of a metallic coating (figure 7). Conversely, the rearward (convex) surfaces exhibited appreciable erosion; being most pronounced toward and along the trailing edges (figure 8). The metallic coating appeared to have appreciably eroded away on this side.

FIGURE 1

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General details
Date: 23 May 2001 Investigation status: Completed 
Time: 1352 hours EST  
Location   (show map):Canberra, Aero. Investigation type: Occurrence Investigation 
State: Australian Capital Territory  
Release date: 10 September 2002 Occurrence class: Technical 
Report status: Final Occurrence category: Incident 
 Highest injury level: None 
 
Aircraft details
Aircraft manufacturer: S.A.A.B. Aircraft Co 
Aircraft model: 340 
Aircraft registration: VH-EKX 
Serial number: 340B-257 
Type of operation: Air Transport Low Capacity 
Damage to aircraft: Minor 
Departure point:Canberra, ACT
Departure time:1350 hours EST
Destination:Sydney, NSW
 
 
 
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Last update 13 May 2014