Failure analysis of two sets of aircraft blades

doi:10.1016/j.engfailanal.2008.10.015

Engineering Failure Analysis

Volume 17, Issue 3, April 2010, Pages 641-647

https://doi.org/10.1016/j.engfailanal.2008.10.015 Get rights and content

Abstract

This paper describes the methodology employed for the failure analysis of aircraft blades and its application to two premature failed sets. The first one corresponds to the high pressure compressor manufactured in a 718 nickel base superalloy. The failure analysis carried out on this blade points towards foreign object damage (FOD). The second set belongs to the high pressure turbine of another engine. Scanning electron microscopy attributes the first fail to the premature failure by a thermo-mechanical fatigue mechanism of one blade with an inadequate microstructure. The remaining blades of this set, which possess a correct microstructure, failed due to the impacts of the debris generated by the fracture of the first one.

Introduction

Modern gas turbine engines for aircraft applications are generally considered to exhibit a high level of reliability, and in-service failures are rare. Nevertheless, this perception is incorrect with a fairly high number of components being retired from service for incipient failure symptoms during the overhaul and maintenance operations. The components most commonly rejected are the blades, from both the compressor and the turbine, and the turbine vanes. The two principal causes are damage caused by ingested materials and the high temperature operation [1]. These two main mechanisms of failure are briefly discussed.

Gas turbine engines can be subjected to ingestion of hard particles or birds, which induce the so called foreign object damage (usually known as FOD). This damage takes the form of sharp V notches in the leading edges of the blades where these foreign substances impact [2]. Dimensions of these notches vary from few micrometers to tens of millimeters, depending on the size, the nature and the severity of impact which induces them. Any gas turbine engine ingests large amounts of air when in operation, either sucked in by the compressor or rammed in by the forward motion of the aircraft. In either event any solid material entrained with the air will cause damage through either erosion or impact [1]. The importance of this failure mechanism cannot be rejected as according to the Boeing web page it causes losses of millions dollars every year to the airlines and airports [3].

Turbine blades operate at very high temperatures, very near of the edge of metallurgical alloy development. This working condition implies, additionally to be required to resist high mechanical loadings, that the material is degraded along the in-service life. Three possible damage mechanisms threaten the integrity of the turbine blades; creep, multiaxial fatigue (associated with the interaction of low cycle fatigue in their longitudinal direction and vibrations induced by the gas flow in the perpendicular one) and high temperature corrosion. Under normal conditions, blades should never be operated at excessive temperatures for long enough periods to cause microstructural damage to the material although exposure for very limited periods of time to more elevated temperature is permitted. Even if FOD is not usual in turbine blades a potential origin of the failure must be found in a very similar mechanism designed as domestic object damage (DOD), which arises from a dislodged debris or component from another location of the engine [4]. It must be remarked that frequently the failure of one component unleashes the fracture of other ones.

The aim of this paper to present the methodology employed for determining the root causes for the failure of the blades and its application to two sets of blades. The first studied samples consisted in a set of blades from the high pressure compressor of an aircraft engine. The second set was constituted by various turbine blades from another engine. Both engines failed prematurely.

Section snippets

Experimental procedure

As indicated above, the methodology employed for analyzing the root causes of the failure of the blades is presented in this paper. The first step consists in a visual examination of the failed blades. This task is performed by the naked eye or with the help of a small stereoscopic microscope (×50) which allows detecting some facets that could have passed unattended. A special attention is paid to the fracture surfaces but without forgetting other aspects which could help identifying the origin

Compressor blades

This set of blades was manufactured using a 718 nickel base superalloy which represents the presently most widely used nickel alloy. This material is strengthened by γ″ particles. Visual examination revealed that most marked damage is sited between the top of the blades and their leading edge. A very marked deformation and even a significant loss of material that was torn away, was observed in this areas. A first analysis of the morphology of this damage points to impact(s) as responsible for

Conclusions

a.
The methodology for determining the root causes of the failure of the blades has been presented and applied to the analysis of the premature failure of two sets of blades.
b.
The whole set of compressor blades failed by the so called foreign object damage (FOD) mechanism due to the impact(s) induced by the ingested sand and stones.
c.
Thermo-mechanical fatigue was blamed as the mechanism responsible for the premature failure of the first turbine blade, fracture being initiated at one of the cooling

References (9)

T.J. Carter
Common failures in gas turbine blades
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E. Silveira et al.
Failure analysis of a set of compressor blades
Eng Failure Anal
(2008)
D. Nowell et al.
Prediction of fatigue performance in gas turbine blades after foreign object damage
Int J Fatigue
(2003)
D. Nowell et al.
Stress analysis of V notches with and without cracks, with application to foreign object damage
J Strain Anal Eng Des
(2003)

There are more references available in the full text version of this article.

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Citation Excerpt :
The fatigue failure of compressor components proceeds in an order of crack initiation, growth and final fracture due to repeated stress concentration after initial damage. The initial damage results from various factors [4–16], such as Foreign Object Damage (F.O.D), corrosion, friction, and notched shape. Fatigue cracks caused by F.O.D can be easily inferred in some cases because apparent damage from impact is observed at the crack initial site, and the chemical composition of impact object can be confirmed at the damage site [5,6].
The flame-out occurred during the ground slide for take-off. As a result of immediate disassembly and inspection after failure, damages to blades and vanes were observed from the stage IV of a compressor. In particular, all stator vanes of stages IV and V, five rotor blades of stage V, and a rotor blade of stage VI were failed. In a fractured stator vane of stage IV, a fatigue-fracture surface, which was originated from the tip, appeared. However, the origin of fatigue cracking was not observed due to severe oxidation and wear. In the case of rotor blades at stages V and VI, the fatigue cracking was initiated from the nick damage part. The same nick damage was also observed in the stator vane of stage V, but fatigue cracking started from the convex surface. Comparing the microstructures of sound and fractured products showed that the convex surface of the fractured product had coarser grains compared to that of the sound product. This indicates that the coarse grains led to premature failure by fatigue cracking.
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^☆: Paper presented at the 17th European Conference on Fracture, Brno, Czech Republic, September 2008.

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Failure analysis of two sets of aircraft blades☆

Abstract

Introduction

Section snippets

Experimental procedure

Compressor blades

Conclusions

Eng Failure Anal

Eng Failure Anal

Int J Fatigue

Stress analysis of V notches with and without cracks, with application to foreign object damage

J Strain Anal Eng Des