Failure as a Design      Criterion

   Fracture Mechanics

   Failure Analaysis

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Wire Rope Failure


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Undercarriage Leg Failure
- Part 1
- Part 2
- Part 3
- Part 4
- Activity 1 - Fractography A
- Activity 2 - Fractography B
- Activity 3 - Fracture Stress


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Aircraft Towbar Failure

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Hail Damage

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Insulator Caps

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Fractography Resource




Dr David J Grieve (d.grieve@plymouth.ac.uk) developed the interactive Java pages for this website.

Background

A Piper PA36-375 Pawnee Brave crop duster aircraft suffered an accident during landing, in which the main landing gear collapsed. The aircraft had flown a total of some 4217 hours at the time of the accident, and had completed about 88 hours since a rebuild. The owner stated that during touchdown, he noticed something passing under the aircraft on the left side. The aircraft then collapsed onto its belly and slid to a halt, causing extensive damage. The failure had involved fracture of both lower saddle brackets and the two main attachment bolts (see Figs. 1 and 2), which allowed the landing gear legs to rotate round and collapse. The left main landing gear was found on the runway some distance from the final position of the aircraft, while the right main leg was located beneath the aircraft.

Each main landing gear leg is in the form of a one-piece steel leaf-spring strut, which is attached to the fuselage frame by an outboard saddle clamp assembly and a single inboard attachment bolt (Figure 1). The saddle clamp assembly comprises a pair of 'L' shaped clamp halves, fitted back-to-back (Figure 2) with a corresponding pair of 'L' shaped plastic pads fitted inside the clamp halves. These prevent fretting contact between the saddle clamps and the spring blade, and assist in uniform spreading of the contact pressure. Saddle clamps are made our of a high strength and high toughness 4340 quenched and tempered steel.  Typical properties and applications for such a steel are hyperlinked from this page, at Mechanical Engineering, Portland State University (http://www.me.pdx.edu/~wernc/eas213F2001/matldata.pdf), at Metal suppliers online, and at Matweb, The Online Materials Information Resource. Specific data for 4340 in the 425oC temper can be accessed from http://www.matweb.com/SpecificMaterial.asp?bassnum=M434AT. Its physical metallurgy is described on the linked pdf file originating from Materials Science and Engineering at Ohio State University.
Crop_Sprayer1.gif (8851 bytes)
Figure 1
Crop_Sprayer2.gif (14755 bytes)
Figure 2

Initial Examination and Basis of Litigation

Examination of the landing gear and saddle components showed that failure was due to fracture of the lower saddle clamps at the change of section from the strap element to the bolt housing (Section F-F in Figure 2), coupled with fracture of the inboard attachment bolts. As is normally the case in aircraft failures, the fracture surfaces were inspected to see if any evidence of fatigue crack growth could be found. The bolt fracture surfaces showed a cup-and-cone appearance (Figure 3). Both of the lower saddle clamp brackets, in contrast, looked significantly different (left hand clamp shown in Figure 4). Significant corrosion damage was present on the fracture surfaces of the saddle clamps, even after light ultrasonic cleaning.  Activity 1 considers the import of these three observations.

Crop_Sprayer4.jpg (23902 bytes) Crop_Sprayer4.jpg (23902 bytes)
Figure 3                        Figure 4

According to the service manual, all main landing components should have been inspected for cracks during the re-build of the aircraft, and replaced if cracked. It was not clear from the manual whether inspection of the clamps required anything more than visual means, although the manual did state clearly that Magnaflux inspection should be performed on the landing gear struts. At best, in the present case, a visual inspection had been performed on the clamps, apparently because of factors like time pressure to return the aircraft to service. Thus the maintenance engineer was held responsible by the aircraft owner and his insurers, because fatigue cracks had unequivocally existed in both saddle clamp brackets at the time of failure and, apparently, had been in existence for some time.

The maintenance engineer (and his insurers) disagreed with this, pointing out that in the course of a 'normal' landing attitude, the saddle clamps should not have been subject to the high rotational loads which had caused the clamp straps to be bent downwards and underneath the rear boss. Such loads, it was believed, could be imposed only by accidental impact of the wheels with the ground prior to final landing, or by a relatively steep nose-down attitude during touchdown. Figure 5 illustrates the effect of small angle changes in landing attitude on the resultant force vector R1, in the absence of drag forces. In all cases R1 passes through the saddle clamp and the applied moment would non-existent (Figure 5c), very small (Figure 5b) or small (Figure 5a). Only when a significant drag force is applied (heavy braking and/or impact with an object on the ground) does the resultant force vector R2 cause a significant moment on the clamp strap. As the aircraft is designed to resist normal braking forces during landing, heavy unintentional impact with a foreign object then appears the most likely cause of any applied moment.

Crop_Sprayer5.gif (20907 bytes)

Figure 5
Damage to the propeller showed no evidence of contact with the ground whilst rotating under power, which discounted the hypothesis of a nose-down impact under power, although it left open the question of a stall-type impact with power off. The underlying argument here is whether these components, in the normal course of events, would have survived until the next scheduled maintenance when the cracks would have been detected.  The argument here is that the plaintiff has flown the aircraft in a way inconsistent with 'normal' use, and therefore bears responsibility for the undercarriage failure, even though the saddle clamps contain cracks.  Although the legal basis for such an argument might appear tenuous, it must be remembered that aircraft components are designed to be 'defect tolerant', and the manual implicitly demonstrates this by specifying crack detection procedures.  Thus the presence of a crack per se, does not necessarily threaten the structural integrity of the undercarriage leg.  Associated with this is argument as to how long the crack had been in existence, which cannot be definitively proven.

Consequently, and as the quantum of loss was significant, the insurers of both the plaintiff and defendant commissioned various reports on the accident from metallurgists and forensic engineers. On the plaintiffs side, these endeavours largely centred around demonstrating the clear existence of fatigue cracks and emphasising the maintenance facility's purported lack of attention to required crack detection. The defendants experts, on the other hand, attempted to calculate failure loads of the cracked saddle clamp, together with flight attitude analyses of the landing loads on the undercarriage legs. These studies were designed to show that a heavy impact load must have caused additional damage to the 4340 quenched and tempered steel saddle clamps for them to fail, even in the presence of the acknowledged fatigue cracking.

At this point, additional evidence from fracture mechanics and fractography was sought to support the defendants case. This was provided by two thrusts; firstly more detailed fractography and, secondly, a first order calculation of the fracture stress on the saddle clamp using linear elastic fracture mechanics.

Proceed to second part of case study.

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