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Monday, March 21, 2005

A310 Loses Rudder, Prompts Fleetwide Inspections, Inquiry

An Air Transat A310 lost its rudder on March 6, prompting manufacturer Airbus to issue a directive calling for field inspections of rudders on both A310 and A300-600 models.

About 400 aircraft worldwide are affected by the inspection mandate. This is a two-fold improvement over the handling when an American Airlines A300-600 lost its tailfin, in that the manufacturer is now calling for immediate inspection of all potentially affected aircraft and using an instrumented tap hammer to do so.

The A310 Rudder Separation ex Cuba

From a pilot's point of view, it is an entirely different proposition to lose a rudder than it is to lose an entire fin and its attached rudder - such as the American Airlines Flight 587 A300-605 did in Nov. 12, 2001, over New York's Jamaica Bay. Loss of a vertical fin would remove all important directional stability, whereas loss of a rudder merely deprives the pilot of most of his ability to control the aircraft in the yaw axis.

"Most," because he can still use the yawing secondary effect of roll and can still turn, climb, and descend almost as normal - as long as both engines are operable and thrust-symmetrical. Once it's clear that the rudder is of use (and essential) only in crosswind takeoffs and landings and one-engine inoperative asymmetric flight, you can appreciate how the Air Transat Flight 961 pilots were able to turn around and land their 1991 built A310-308 at Varadero, Cuba.

Nevertheless, the event has refocused attention on A300 and A310 rear ends and stimulated attempts to relate this incident to the AA587 accident. The fact that the A300 and A310 have identical composite (carbon fiber) fin construction, rudder actuators and software does nothing to dispel the notion of a possible causal relationship. AA587's pilot was blamed by the investigators for pedaling his aircraft's rudder pedals into an involuntary aircraft/pilot coupling (APC) during a wake turbulence encounter (ASW, Jan. 31). He allegedly overstressed the vertical fin to as much as 1.93 times its design ultimate load. By comparison, the Flight 961 pilots were cruising at Flight Level 350 on autopilot when there was a sideways lurch that threw a flight attendant off her feet, a loud bang and a change in handling characteristics. Naturally, the pilots weren't aware until after they'd landed that 95 percent of their composite construction rudder had departed.

Airbus Operator Information Telex TX530526F of March 8 advised that "a portion of the rudder structure remained attached between the lower hinge and the three servo-control attachment points." The big difference between the American Airlines event and Air Transat one may be related to the A310's rudder having momentarily aerodynamically loaded up an unstressed and streamlined fin (before itself breaking) and not a fin already side-loaded by the combination of repeated out-of-phase pilot pedal inputs reinforcing a yawing cycle that was first initiated by wake turbulence. Consider also that American Flight 587 involved a 266-seat A300's fuselage length of 177'5" (54.1m) compared to a 200-seat A310's length of 153'1" (46.66m). The A300 has a 27-foot-tall fin and the A310 has a slightly shorter fin. The difference in length between the two aircraft would also vary the moment arm for stability-restoring forces acting upon the vertical fin.

Composite Past and Future

The A310 series featured the first composite primary structure to appear on non-military aircraft. The original A300s had aluminum fins. Currently Boeing's 777 and all in-production Airbuses have composite vertical and horizontal stabilizers. The Boeing 787 will be all composite and the A380 will have an all composite center wing box, aft pressure bulkhead, tail surfaces and many other composite parts. In the bizjet field, Raytheon's Premier I and Hawker 4000 Horizon both have entire fuselages of carbon fiber/epoxy honeycomb, while Visionaire's VA-10A Vantage promises to be the world's first all composite business jet.

But it's Boeing and Airbus that have a definite need to resolve this failure mechanism. If the composite tails are starting to age via a wear and tear process, it would seem to be more likely that freezing and refreezing of trapped water or moisture entrained in the matrix voids or between laminations would be a likely culprit. These aircraft move between destinations at very cold and very hot latitudes and regularly spend much time at high-level cruise in outside air temperatures down to minus 50 degrees C (-58�F). In the AA587 case, the rudder broke itself into three pieces and departed the vertical fin at some stage. An early AA587 theory suggested that the rudder may have fluttered (very rapid fullish movements), failed at the bottom hinge and then rotated laterally around the top hinge before tearing away with the overstressed fin. That theory faded into the background as being the outcome, a symptom of a catalyst triggering mechanism that had to be either the flight-control system or involuntary and oscillatory cyclic inputs via the pilot's rudder pedals. These latter inputs are referred to as reversals. If they become out-of-phase with the aircraft's yawing, they reinforce the yaw-induced loads upon the fin and can easily induce fin attachment failure. That fact was almost unknown before it was disclosed by the AA587 investigation. As few as three sequenced opposite rudder inputs (known as a triplet) could in fact overstress the fin.

Concorde's Rudder Failures

British Airways' fleet of seven Concordes was re-equipped with new rudders in 1992, following a series of rudder problems. In total there were six cases of partial rudder loss. All were attributed to defects brought to light by fatigue-induced failure. Their construction was metal honeycomb bonded to an aluminum alloy skin with a film adhesive. None of these defects were detected prior to failure, and all departed with a vibration and a final airframe-transmitted "thump." Significantly, the trailing edge was "unfastened" so the trailing edge bond's "disbond" just propagated forward into the core-to-skin bond. Skins separated from their honeycomb core must flutter and accelerate the separation disbonding.

Composite Fatigue?

The black art of scientific composite inspection can use thermal wave imaging and ultrasonics, but in service it usually relies upon visual examination and aural tap-tests ("does it sound right?"). Researchers at the U.S. Air Force Research Laboratory's Materials and Manufacturing Directorate comment that "substandard fabrication procedures, environmental exposure and handling, or service damage can all have a negative impact on the mechanical integrity of these structures without affecting their visual appearance."

Failure modes are very complex. Unlike metal fatigue, which is about single crack growth (often branching into secondary ones), composites degrade through tiny cracks within the matrix. This reduces the stiffness and ultimately the loads for failure. Also, any blunt impact (due to ramp-rash, say) will start cracking, leading to what's known as "BVID" or Barely Visible Impact Damage. BVID will cause a step reduction in material strength in the impact area. Frankly, nobody understands these well, for which reason all the airworthiness codes require much higher strength from composite components than they do from metal ones -- usually 25 percent to 50 percent more (referred to in certification circles as "the composite superfactor").

When a composite component fails, it was thought to be unlike a metal. i.e., it keeps appearing perfectly intact right up to the point of failure, then goes almost instantly -- not like metal parts which start developing bends. If you remember high school physics terminology, the Hookes Law period supposedly extends right up to the failure point, there being no plastic deformation. With later lay-ups, that was proved wrong. A composite material, which is intrinsically anisotropic, does not fail like isotropic metals, but it can fail progressively in a non-catastrophic fatigue-like mode. Composites also follow the linear elastic behavior until a certain stress and elongation, where tiny micro-cracks start to form in the matrix material (e.g., an amine cured epoxy resin like the MY720 resin from Ciba/Vantico with a DDS hardener). These micro-cracks start to form larger cracks within the composite but the component is still far from catastrophic failure.

What happens? The incorporated fibers (e.g., carbon fibers) act initially as crack-stoppers and further progression of the crack needs more energy and therefore more stress. The Young's Modulus of the composite is now lower due to the fact that new "inner surfaces" are generated by the cracks. Further loading the composite will lead to the initial failure of individual fiber filaments and yarns with toughening effects like fiber bridging, disbonding from the matrix material and finally very important fiber pull-out effects.

These effects give the composite a "quasi-plastic" failure behavior leading to non-catastrophic failure. Indeed, this depends on the fiber-layout within the composite, i.e., unidirectional layout compared to a 0/90 or 0/45/90/45/0 layout. Composite science is a very complex area and a lot of work has been done in the past 20 years to understand the micromechanical aspects of damage, both static and/or dynamic. Reliability of composites, metals and especially ceramics is governed by the laws of crack growth covered in the science and technology of fracture mechanics, again both static and dynamic. Understanding these laws and applying them correctly, combined with the knowledge of environmental aging, is the basis for their use in the aviation industry.

Known Failure Modes

When an aircraft changes from low temperature, high altitude operation (cruise) to a low level, low altitude condition (landing), the low pressure inside the composite material with microscopic cracking will draw in air, which will accumulate water inside due to condensation created by the cold interior. With a return to high altitude, low temperature operation, the water will freeze inside and NOT be expelled on landing even after the water returns to a liquid state. With repeated cycles of low pressure, low temperature operation to areas of high humidity and high pressure, water can accumulate inside a composite material, which has voids in which air can enter. Eventually, water will accumulate in these voids until its freezing will cause material separations (loss of strength) within a composite assembly. Enough of these microscopic (internal and unseen) separations and operational cycles may lead to complete failure of the composite assembly during "normal" operation.

"Airbus will be taking an abundance of caution," company spokesperson Clay McConnell said, after disclosing that more than 400 A310 and A300-600 aircraft worldwide will be inspected over the next few weeks. "A310 and A300-600s have identical rudders so there will be visual inspections and tap aural tests carried out to detect any hidden flaws."

Marc Fernandez, senior investigator with Canada's Transportation Safety Board, said that the TSB will be removing the fin off the incident A310 in Cuba and shipping it to Airbus in Germany.

BA Concorde Rudder Failures

  • Nov. 27, 2002: New York flight loses part of lower rudder.
  • Oct. 8, 1998: Part of lower rudder detaches off Newfoundland.
  • March 21, 1992: Large section of upper rudder lost at twice speed of sound.
  • Jan. 4, 1991: Portion of upper rudder separates during London -- New York flight.
  • April 12, 1989: Part of upper rudder detaches at 44,000 feet.

Source: UK AAIB

Hookes Law

The principle that the stress applied to stretch or compress a body is proportional to the strain or to the change in length thus produced, so long as the limit of elasticity of the body is not exceeded.

Source: http://www.answers.com/topic/hooke-s-law

Young's Modulus

Within the limits of elasticity, the ratio of the linear stress to the linear strain is termed the modulus of elasticity, or Young's Modulus, and may be written Young's Modulus, or E = (Stress/Strain). It is this property that determines how much a bar will sag under its own weight or under a loading when used as a beam within its limit of proportionality.

Source: metals.about.com/od/metalterminology/l/bldefyoungsmodu.htm

 

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