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Further Reading

F-111 Problems

The F-111 introduced concurrency (overlap) between design validation/verification and production to accelerate program schedules because there was an urgent need for the weapon system capability; however, technical problems uncovered after the design was "frozen", and production was underway, led to costly retrofits and schedule delays. The F-111, like any complex weapon system development program which provides new war-fighting capability, had areas of risk or deficiency that came to light during RDT&E even though there was perceived low risk in the design. Verification and validation testing of the design is meant to both verify performance and uncover unanticipated design shortfalls. Ideally, the verification and validation leads production so that if changes are necessary, there is no impact on production, and a minimal impact on program cost.

However, technical problems uncovered during F-111 validation and verification led to costly retrofits and redesign of the production versions. Unlike the transonic drag risk area, which was known before F-111 RDT&E began, there were other risk areas which came to light during RDT&E - in the case of the F-111 - after Critical Design Review and first flight. A catastrophic failure in the wing pivot structure as late as 1969 caused the entire F-111 fleet to be grounded for several months, and required cold-proof testing of all models which had been built to that time and subsequent production units. Both of these risk areas required excessive resources by the contractor and the F-111 System Program Office to solve, with significant retrofits of over 200 previously-built aircraft, and a schedule impact on F-111 production and operation date".

The F-111 aircraft was conceived as a low-level flight supersonic strike aircraft capable of terrain-following flight in all weather and visibility conditions. This in turn required some unique characteristics of the design, such as the variable sweep wings, which would provide the aerodynamic performance for low-altitude supersonic flight while allowing for take off and landing with a very high weapon load. The design of the aircraft meant that all wing loads were transferred into the fuselage through the wing pivot mechanism. Furthermore, with the wing at a fully swept back position, the elevators in the empennage control not only the pitch of the aircraft, but also its roll. This means that rear fuselage of the aircraft has to withstand twisting as well as bending loads.

The whole concept of the aircraft called for high strength materials in the airframe with high strength/weight and strength/volume ratios in the critical structural areas. Early on in the design stage, it was decided to utilise ultra-high strength (UHS) steel for the structure-critical components, since it had better strength to weight properties than the aluminium alloys used traditionally (until then). The steel that was selected was a medium-carbon low alloy Ladish Corporation steel designated D6ac.

The F-111 did not have particularly happy start to its service life, due to several structural failures both in-flight and during ground fatigue testing. The cause of the failures was ultimately attributed to a large variation in fracture toughness of the D6ac steel, with the lower limit of the toughness values being unacceptably low. The initial defects were created, almost universally, during manufacture of the steel components. The low toughness of some components or even in some specific locations on individual components meant that only relatively short fatigue crack growth had to occur before the crack reached catastrophic length. However, as the length of service of the F-111s increased, the failures originating from manufacturing flaws were replaced by failures caused by degradation of the material in service.

A fatigue test of the full aircraft was started in Fort Worth, Texas in August 1968. Test article A4 failed after just 400 hours of testing, foreshadowing the inflight failure with uncanny accuracy. The failure originated from a bolthole in the aft surface of the WCTB near the junction with the bottom plate. The failure was traced to poor manufacturing processes. With the start of the next test, test article FW-1, developed a critical crack after 2800 hours. While this was a grand improvement over 400 hours, it was by no means adequate. Trouble continued with article FW-2. In June 1969, at 7800 hours, it suffered a catastrophic failure in the outboard closure bulkhead.

The first documented in-flight failure for the F-111 was early in 1968. In this accident, an F-111A crashed during deployment in Vietnam due to a sudden catastrophic failure in the tailplane system. The origin of the failure was traced to a fatigue fracture of a welded joint in the power unit of the left tailplane [Gunston 1987]. This misfortune was repeated on 8 May 1968 when another aircraft was lost near Nellis AFB in the US for exactly the same reason. Unfortunately, it is not clear whether this component was made from D6ac steel.

The most familiar of all F-111 in-flight failures involved aircraft 67-049. Reaction to this accident was widespread throughout the US Air Force, the airframe contractor, and subcontractors alike. Fallout from the loss of this one aircraft shaped several programs in flight safety that continued including, for the F-111 specifically, and for USAF aircraft in general, the adoption of the damage tolerance design philosophy. The accident aircraft was an F-111A, which had accumulated just over 100 flight hours; it crashed on the Nellis AFB range on 22 December 1969. During pull-up from a rocket-firing pass, a fatigue crack in the wing pivot fitting reached a catastrophic length, and the left wing separated from the aircraft. The crack formed in the 7.26 mm thick lower plate of the WPF from an initial manufacturing flaw 5.72 mm deep. The crack then grew a mere 0.44 mm by fatigue to a critical depth of 6.16 mm during the span of 104.6 flight hours. At the time of failure, the crack had a total surface length of 23.6 mm. Of great concern in this catastrophic failure was both the very small depth of the critical flaw and the extremely short time of fatigue crack growth.

The failure on this aircraft resulted in the fleet being grounded. Aircraft were released for flight after being subjected to a process known as the Recovery Program, and it involved testing at -40°C under two load conditions, -2.4 g and +7.33 g, at 56 degrees of wing sweep.

Two other aircraft experienced failures under the Recovery Program. The first was an F-111E in 1970 [Laffe and Sutherland 1994]. The left hand horizontal tail pivot shaft failed at 88% of the maximum load applied in the Cold Proof Load Test (CPLT) at the Fort Worth test facility. Shortly thereafter in 1971, another Recovery Program aircraft suffered catastrophic failure. This failure occurred in an F-111A at the Sacramento Air Logistics Center (SM-ALC). The crack completely ruptured the lower plate along with the front and rear spars.