Madan, M and Sujata, M and Raghavendra, K and Bhaumik, SK (2011) Analysis of components of aeroengine retrieved from the wreckage of an Aircraft. Project Report. National Aerospace Laboratories, Bangalore, India.Full text not available from this repository.
A few components belonging to an engine retrieved from the wreckage of crashed aircraft were sent to this laboratory for investigation. The components that were sent for laboratory analysis are (i) fractured rim and diaphragm of stage IV compressor disc, (ii) spline bolt, (iii) HPTR blade, (iv) LPTR blade, (v) HP nozzle guide vane, (vi) stage I compressor rotor blades, and (vii) a metallic piece recovered from fuel oil unit. Examination revealed that the diaphragm of stage IV compressor disc had sheared off circumferentially and along the fillet of the rim. While majority of the fracture surface showed reverse slant fracture features, a region spanning about 50-60 mm on the fracture surface had smooth and flat appearance. Fractographic study conducted on this region showed presence of a half-moon shape region, indicative of progressive mode of crack propagation. Further examination of this region under a scanning electron microscope confirmed that the crack propagation was by fatigue. The fatigue crack was found to have initiated at the fillet of the rim of the disc on the side facing stage V compressor. Detailed examination revealed that after initiation, the crack had propagated progressively over about 18 mm in length and 2.5 mm in depth. Followed by this, the fatigue crack propagation was relatively fast through the remaining thickness of the diaphragm and over a length of 50-60 mm on the diaphragm surface. Once the fatigue crack attained this length, the diaphragm sheared off circumferentially and along the fillet of the rim instantaneously by overload. The fatigue crack origin region was severely damaged due to post fracture rubbing and hence, no further study was possible to examine whether or not there was any localized stress concentrator either mechanical or metallurgical, which could be responsible for the fatigue crack initiation. It may, however, be noted that there were no metallurgical abnormalities in the material of construction of the compressor disc and it was found to conform to specification of BT3-1 Ti-alloy. Study of the surfaces and the metallurgical examination did not reveal any evidences of overheating in the components such as HPTR blade, LPTR blade and HP nozzle guide vane. Although certain microstructural changes were observed in the material of HPTR blade, these changes were found to be within the acceptable range, and they are expected to occur in the material over prolonged use at elevated temperatures. Although the features on the fracture surfaces of the spline bolts were completely obliterated due to oxidation/exposure to fire, gross fractographic features and deformation preceding fracture indicated that the failure was by overload. Compositional analysis confirmed that the metallic piece recovered from the fuel oil unit is a fractured segment of the nose-fairing. Analysis also showed presence of traces of nose-fairing on the damaged regions of stage I compressor rotor blades. Analysis suggests that the compressor disc with a fatigue crack of length 50-60 mm at the rim fillet region would not survive, and it is likely to fracture during engine run. It is, therefore, stated that the fatigue fracture of stage IV compressor disc was the first in the chain of events that led to engine failure.
|Item Type:||Proj.Doc/Technical Report (Project Report)|
|Uncontrolled Keywords:||Stage IV compressor disc;Fatigue failure|
|Subjects:||AERONAUTICS > Aeronautics (General)|
CHEMISTRY AND MATERIALS > Chemistry and Materials (General)
|Division/Department:||Materials Science Division, Materials Science Division, Materials Science Division, Materials Science Division|
|Depositing User:||Ms. Alphones Mary|
|Date Deposited:||07 Jul 2011 14:48|
|Last Modified:||07 Jul 2011 14:48|
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