At present many aircraft that are approaching the calendar-based limit of service life (defined by manufacturers) are stili in operation. Most of them show considerable reserves of fatigue life due to small numbers of flying time consumed. Further operational use will be possible afler completion of a large-scale testing project with various non-destructive methods engaged. Therefore, knowledge of changes that have occurred on the surfaces of structural members of an airframe and of engine components comes into prominence. Most dangerous to operational use of products of aeronautical engineering are fatigue cracks that nucleate at locations of no direct access to critical areas. The critical areas of an engine structure include surfaces and edges of compressor and turbine blades.
A crack of the first-stage compressor rotor blade occurred in the course of operating one of the aircraft types. The crack was of fatigue nature and occurred in only a slight distance from the blade bottom surface fixed in the blade locking piece. Such a crack is extremely dangerous because it can result in an aircraft crash. The problem that faced the flaw-detection diagnostics resolved itself into applying such a flaw detection method that would help finding the nucleating fatigue cracks in a reliable way, with aircraft availability not impaired. Both the failure location and no direct access to this area made that ultrasonic testing was recognized the only flaw detection method useful in that case.
Digital flaw detector of the SOFRATEST type with special probe heads were developed to serve the purpose during laboratory examinations. Original compressor rotor blades with artificial standard cuts served as reference reflectors. Minimal dimensions of such a cut were determined in the course of metallographic examination of a real fracture.
A reference failure was originated in the blade locking piece, at a distance of 7 mm from the bottom surface of the blade locking piece. The reference reflector was 0.3 mm deep and 5 mm long and was located in the middle of the blade, on its upper surface. Small dimensions of the reference reflector and complicated geometrical forms of the blades under examination forced the 'flaw detector - probe head' system to operate at high gain level.
In the course of examining the blades some difficulties appeared, mainly in distinguishing the apparent echoes from the shape- and failure-originated ones. Difficulties that were arising in the course of examining the blades were as follows:
a) fading of a final pulse in the cases of:
b) occurrences of echoes from: |
As the pulses originated by a fatigue crack and an apparent failure are represented on the flaw detector display very close to each other, the essential diagnostic problem was to take the right decision.
A suggestion was put forward on fundamental differences in variations of pulses originated by a fatigue crack as compared to those originated by apparent failures. The hypothesis was based upon the assumption that there is a fundamental difference between the nature of a fatigue crack and that of apparent failures. Pulse spectrum being the effect of applying the Fast Fourier Transform (FFT) was accepted as the objective measure of recording the difference in reflected pulse variations. The analysis showed differences in spectra of signals originated by apparent failures and those of signals originated by failures of the fatigue-crack type. Interpretation capabilities have therefore been considerably extended and possibilities for identification of echoes - provided.
By way of example, Figs. 1.1 - 1.4 show some recorded pulses as reflected by the failures, the Fourier transforms, i.e. the spectra - included. The spectrum of a pulse reflected by the reference failure (Fig. 1.2) includes two envelopes, i.e. the main one with the midband frequency of 3.750 MHz, and the side one with frequency of approximately 11 MHz. The bandwidth under the main envelope is 12.366 %.
Fig. 1.1. The ultrasonic signal as recorder with the digital flaw detector High pulse of a failure (a reference cut) on the blade locking piece is evident.
Fig. 1.2. The Fast Fourier Transform (FFT) of the ultrasonic pulse as reflected from
Fig. 1.3. The low ultrasonic signal from the apparent failure precedes the high ultrasonic signal originated by the cut.
Fig. 1.4. The Fast Fourier Transform (FFT) of the ultrasonic pulse as reflected from the apparent failure.
The spectrum of a pulse reflected from the apparent failure (Fig. 1.4) consists of one main envelope with the midband frequency of 3.87 MHz. The bandwidth under the main envelope is 23.534 %. There is a distinct difference between both the represented spectra. The analysis shows that the spectral bandwidth is a diagnostic parameter that discriminates the apparent failure from that of the fatigue-crack type. Experience gained while examining the first-stage blades of the aero-engine compressor with the SOFRATEST digital flaw detector enabled the ultrasonic blade inspection procedures to be developed and implemented.
The Polish-made flaw detectors UNIPAN 520 and ultrasonic probe heads of the 4S7 type (fitted for examining critical areas of the blade) proved to be useful with this method. Original compressor rotor blades with reference cuts were used as standard items.
Increase in amplitude of the pulse reflected from the apparent failure was observed, i.e. the phenomenon effected by intense moistening of the engine inlet duct. Increase in the effective reflecting surface of the apparent failure was the reason for the phenomenon to occur. The ultrasonic inspection of blades was recommended to follow the engine test to avoid pulses originated by the apparent failure to appear on the flaw-detector display. In the course of engine running the points of contact of blade locking pieces with disk grooves are heated and dried. The effective reflecting surface is therefore kept to a minimum.
Author wishes to thank Mrs. Malgorzata MERCIK for the English form of the paper.
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