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.
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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.
As F-111 operations expanded within the Air Force in the late 1960's, a rash of incidents involving unexpected departures from controlled flight during maneuvers at high angles of attack occurred. The F-111 had been designed with a g-command flight control system that provided g-forces in direct pro-portion to the deflection of the pilot control stick. However, in providing the pilot with the level of g-force, the system would increase the angle of attack of the aircraft. Unless the pilot was monitoring the angle of attack, the aircraft could enter a range of high angles of attack where a loss of directional stability resulted in an unintentional yaw departure and spin entry. These findings led to an Air Force program in 1973 to develop a stall inhibitor system (SIS) for the F-111. Several F-111 aircraft were lost in spin accidents during fleet operations; however, the subsequent implementation of the SIS prevented stalls and eliminated spins as an operational concern.The F-111 airframe utilized a significant amount of high-strength D6ac steel in the wing carry-through structure. This component was heat treated to a tensile strength of 220,000 psi and designed for -3g to 7.33g with design flight life goals of 4,000 hr and 10 years of service. However, a full-scale static test program that was conducted over a 6-year period encountered several failures, including a failure at the wing-pivot fitting. Various modifications, including the first use of an advanced boron-reinforced composite doubler to reduce stress levels, coupled with an extension of the structural tests to 40,000 hr, were believed to have provided for 10,000 hr of safe operations. In December 1969, an F-111 experienced a catastrophic wing failure during a pull-up from a simulated bombing run at Nellis Air Force Base. This aircraft only had about 100 hr of flight time when the wing failed. The failure originated from a fatigue crack, which had emanated from a sharp-edged forging defect in the wing-pivot fitting. As a result of the accident, the Air Force convened several special committees to investigate the failure and recommend a recovery program. The original material had low fracture toughness due to the heat-treatment process. The heat-treatment process was corrected to provide improved toughness for the D6ac material.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.
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.
The Sealed Chest is the hardcore version of the Locked Chest, as it doesn't even have a padlock to open it with. Your only option here is to crack them open with explosive force. This means using Rockets, Superflame, Supercharge, or other high-power methods like turrets.
The data collected point to two possible failure routes - cracking of the hull plates or failure of the riveted seams. However, the fact that the hull was not severely deformed, as evidenced by the sonar images and reported by the survivor Fireman Barrett, implies a fairly low energy failure. A suggestion of brittle fracture of the steel plates at ice brine temperatures was made by two groups in Canada in 1991 based on Charpy tests of a hull plate fragment. However, slow bend testing, a more likely applicable strain rate, of four hull plate samples showed average toughness of 55 MPa-m1/2 at 0C, quite reasonable for this application. If it was not brittle steel, were the riveted seams strong enough?
Fatigue cracking is most often related to the age of a metal component and happens when a small crack grows to the point that the cross-section of a component is reduced to the point where it can't carry the load imposed. 2ff7e9595c
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