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AA 587 -- Vertical stabilizer & composites (thread#3)

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AA 587 -- Vertical stabilizer & composites (thread#3)

Old 18th Nov 2001, 04:31
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Post AA 587 -- Vertical stabilizer & composites (thread#3)

Having reached the 100 posts-per-thread cap, CaptPPrune has closed the two main previous threads on this accident. (Titles were new clues deepen AA 587 crash mystery ... and vertical stabilizer AA 587)

Uncertainties and skepticism were voiced on the previous threads about the ability to visually detect whether a composite has deteriorated because of age, stress, or from environmental factors. Might this not suggest the need to do some destructive testing of the composite area in the vertical stabilizer to see whether AA 587 was a one-of, or whether there are other incipient failures lurking in the fleet?

Are there any A-300's or A-310's parked in the desert or elsewhere whose vertical stab composite might be sacrificed to gauge whether the design strength still is present?
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Old 18th Nov 2001, 15:10
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This may be worth noting...

NTSB Identification: DCA99IA058

Scheduled 14 CFRPart 121 operation of Air Carrier AMERICAN AIRLINES
Incident occurred Tuesday, May 11, 1999 at MIAMI, FL
Aircraft:Airbus 300-600, registration: N7082A
Injuries: 129 Uninjured.

This is preliminary information, subject to change, and may contain errors. Any errors in this report will be corrected when the final report has been completed.

On Tuesday, May 11, 1999, at 1445 EDT, an American Airlines A300, flight 916, a scheduled passenger service from Bogota, Columbia, to Miami, Florida, landed successfully after the flightcrew experienced multiple rudder deflections that caused the airplane to yaw excessively from side to side while on final approach to runway 9R. There were no reported injuries to the 119 passengers or crew of 10.

Preliminary information from both the American Airlines engineering group and a Safety Board crew interview indicates that during the initial approach to runway 9R, as the crew configured the airplane for landing with flaps 40 degrees and the landing gear down, the airplane began to yaw left and right. The flightcrew stated in an interview that the rudder pedals in the cockpit did not move, though the rudder was deflecting and causing a yaw motion that was sufficient to prompt the captain to abandon the first landing. During the go around, and specifically, as the airplane was reconfigured with the landing gear up and flaps at 20 degrees for the go around, the yaw deviations increased and became extreme. The crew reconfigured the airplane twice during the go around and completed the landing with 20 degrees of flaps.

Initial information from American's FDR readout indicates that the rudder which is a single panel with three hydraulic actuators, deviated continuously but not rhythmically between 5 and 11 degrees each side of center during both approaches. The FDR has been transported to the Safety Board for additional readout.

[ 18 November 2001: Message edited by: Flight Safety ]
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Old 18th Nov 2001, 16:19
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Just scanned previous thread and saw that 1 post referred to loss of hydraulics with a departed vertical stab. Yes, the Blue and Yellow systems would have been immediately lost as the plumbing ripped away, but the Green system is protected from inflight collision scenarios by a hydraulic "fuse" that should shut in the event of excessive flow rate to protect remaining green fluid from loss. The location however of the fuse would in this case be critical. If it is located anywhere in the vert stab, or near its attach points, then it probably would have gone as well - and there goes your protection, as well as the remaining system, and all your primary flight controls (rudd, ail, elev). Don't have accurate enough documentation to determine exact location of this fuse, but suspect this may have been what happened. In this event, the aircraft would indeed have been extremely difficult to control, even if the crew had immediately resorted to a UAL Sioux City mode of control, which in the speed/power range they were in (climb out) they may not have been able to manage.
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Old 18th Nov 2001, 20:43
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There was a quote on previous thread that Boeing and Douglas do not use composite primary structure , got this form a google search :
Shorts awarded Boeing contract to produce Boeing 737 all-composite rudders. Shorts
offered for sale by the United Kingdom Government (1988) and acquired by bombardier .

also picked up story from flug revue:
http://www.flug-revue.rotor.com/FRHe...07/FR9807b.htm

Composites were at first chosen mainly for weight savings, fatigue and corrosion resistance, but a major new criteria has appeared: manufacturing cost reduction. Initial employment of larger, more demanding primary composite structures at Airbus began in Germany with the single piece rudder for production model A300s and A310s as far back as 1983. The relatively straightfoward design utilized three carbon/epoxy skinned honeycomb sandwich panels assembled as a hollow triangle to form the rudder. About eight meters in lenght and varying from one to two meters in chord, weight of the composite version is 175 kilogramms. It could be directlysubstituted for the metal version without changes to the plane. Improvements over the original metal structure included weight savings of some 22 percent, component count reduced by nearly half plus reduction of details froom over 17000 down to only 4800.

Hardly two years later, the much larger, heavier and far more complex vertical stabilizer was converted from metal to composite. Essentially of carbon/epoxy laminate, with some glass and honeycomb core on the leading edges, and weighing 800 kilogramms, it was for several years the largest primary composite structure flying on a production model civil transport. As with the rudder, design and manufacture is by the German partner DASA Airbus, at its highly automated Stade plant south of Hamburg, a facility specializing in composite structures. The program entailed a running re-design and replacement of the original all-metal vertical stabilizer in all Airbus production models, beginning with the A300.

Two seperate manufacturing methods, matten and module, were investigated for this program, which the latter won. Module technology is a multi-step, semi-automated process whereby aluminium modules are wrapped with prepreg composite material to form integral skin stiffeners. They are then positioned by robot on the already laid up skin, where US strips are in turn placed on top of the modules to form spar caps. Most of the work at Stade except the actual module wrapping has been automated. Even prepreg is carried from the cutters to the layup area by robot vehicles employing carbon/epoxy containers or baskets made from scrap prepreg. The entire, single piece vertical stabilizer skin was then co-cured in a single step, in a large autoclave. The advantage: only one trip to the autoclave was necessary. the relatively massive attach fittings at the base of the fin are also of carbon/epoxy and integrated into the skin; only the pins and bushings are metal.

The original co-curing has since been modified to simplify the tooling, reduce manufacturing risk and cost, however more trips to the autoclave are required. Moreover the massive attach fittings, normally made up of hundreds of layers of prepreg, are being considered for manufacture by another method: Resin Transfer Molding RTM.
 
Old 18th Nov 2001, 23:18
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DEJA VU- Thereby hangs a Tail

Reference this A320 series AD (below):
I would guess that a disbonding over a large area of the vertical stabilizer would make the fin and rudder much more prone to a destructive flutter-mode not encountered in pre-certification testing. After reviewing the comment regarding A330 (also) at the end of this AD, it must be said that there would seem to be few Airbus types unaffected.

____________________________________________
[4910-13-U]

DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 39 [65 FR 37029 6/13/2000]

[Docket No. 2000-NM-139-AD; Amendment 39-11776; AD 2000-11-27]

RIN 2120-AA64

Airworthiness Directives; Airbus Model A319, A320, and A321
Series Airplanes

AGENCY: Federal Aviation Administration, DOT.

ACTION: Final rule; request for comments.

SUMMARY: This amendment adopts a new airworthiness directive (AD)
that is applicable to certain Airbus Model A319, A320, and A321
series airplanes. This action requires a one-time ultrasonic
inspection to detect disbonding of the skin attachments at the
stringers and spars of the vertical stabilizer, and repair, if
necessary. This action is necessary to detect and correct
disbonding of the vertical stabilizer structure, which could
result in reduced structural integrity of the spar boxes of the
vertical stabilizer.


DATES: Effective June 28, 2000.

The incorporation by reference of certain publications listed in
the regulations is approved by the Director of the Federal
Register as of June 28, 2000.

Comments for inclusion in the Rules Docket must be received on or
before July 13, 2000.

ADDRESSES: Submit comments in triplicate to the Federal Aviation
Administration (FAA), Transport Airplane Directorate, ANM-114,
Attention: Rules Docket No. 2000-NM-139-AD, 1601 Lind Avenue,
SW., Renton, Washington 98055-4056. Comments may be inspected at
this location between 9:00 a.m. and 3:00 p.m., Monday through
Friday, except Federal holidays. Comments may also be sent via
the Internet using the following address: 9-anm-
[email protected]. Comments sent via the Internet must contain
"Docket No. 2000-NM-139-AD" in the subject line and need not be
submitted in triplicate.

The service information referenced in this AD may be obtained
from Airbus Industrie, 1 Rond Point Maurice Bellonte, 31707
Blagnac Cedex, France. This information may be examined at the
FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW.,
Renton, Washington; or at the Office of the Federal Register, 800
North Capitol Street, NW., suite 700, Washington, DC.

FOR FURTHER INFORMATION CONTACT: Norman B. Martenson, Manager,
International Branch, ANM-116, FAA, Transport Airplane
Directorate, 1601 Lind Avenue, SW., Renton, Washington 98055-
4056; telephone (425) 227-2110; fax (425) 227-1149.

SUPPLEMENTARY INFORMATION: The Direction Générale de l'Aviation
Civile (DGAC), which is the airworthiness authority for France,
notified the FAA that an unsafe condition may exist on certain
Airbus Model A319, A320, and A321 series airplanes. The DGAC
advises that localized disbonding has been detected on the skin
attachments at the stringers and spars of the spar boxes of the
vertical stabilizer. During the manufacturing process, pre-cured
parts (attachments of the stringers, spars, and ribs) are
installed on the skin panel before the final curing process. A
peel ply is used to protect the contact surfaces of the
attachment angles of the skin panels of the vertical stabilizer
until the pre-cured parts are ready for installation.

Investigation revealed that, after the peel ply was removed from
the attachment angles, a residue of polymer finish contaminated
the contact surfaces of some pre-cured parts.
This contamination
reduced the adhesive strength of the bond and, in some cases,
caused debonding (disbonding) of the skin attachments. This
condition, if not detected and corrected, could result in reduced
structural integrity of the spar boxes of the vertical
stabilizer.


Explanation of Relevant Service Information

The manufacturer has issued Airbus Service Bulletin A320-55A1027,
dated May 12, 2000, which describes procedures for a one-time
ultrasonic inspection to detect disbonding of the skin
attachments at the stringers and spars of the vertical
stabilizer, left- and right-hand sides, and repair, if necessary.
If any disbonding (damage) is detected and the area of damage is
greater than 300 square millimeters (mm2), or if multiple damage
is detected on one specific component (stringer/spar attachment),
the repair involves installing additional fasteners in the
affected areas. The amount of damage determines the number of
additional fasteners to be installed in the affected area.


Additionally, Airbus Service Bulletin A320-55A1027 references
Airbus Service Bulletin A320-55-1026, Revision 01, dated May 20,
1999, which, for certain airplanes, describes procedures for
prior or concurrent modification of the vertical stabilizer to
ensure proper reinforcement of the structure/skin attachments.

The DGAC classified Airbus Service Bulletin A320-55-A1027 as
mandatory and issued French airworthiness directive T2000-208-
148(B) R1, dated May 17, 2000, in order to assure the continued
airworthiness of these airplanes in France.
_________________________________________
From an associate:

"You are aware that there was a big scandal in France when it was discovered that AI had the blank DGAC certification forms already signed and stamped in a drawer at their Toulouse headquarters. I believe some stuff was found to be dealing with "new material specification and certification".
"There was a serious problem at SR last year, when they discovered that their rudders in the brand new A 330 came apart by delamination..."
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Old 19th Nov 2001, 00:40
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Belgique: Interesting Info. Reinforces my intent to avoid being a pax aboard any Bus until composite failures are understood and corrected.
Photos of separated vert stab attach fittings Click on individual photos to enlarge for full screen viewing!

[ 18 November 2001: Message edited by: GlueBall ]

[ 18 November 2001: Message edited by: GlueBall ]
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Old 19th Nov 2001, 02:15
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Belgigue.....
Thanks for your outstanding informatory contributions on this page....much appreciated.....measured......non-biased and objective,,,,,Do you think we are looking at a Ford Pinto scenario here ???....
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Old 19th Nov 2001, 02:42
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More information excerpted from the Nov. 18 Washington Post.

"..the fin sheared off above the brackets at its base, neither the brackets nor the bolts anchoring them to the plane's aluminum
fuselage appear to have failed. The joint itself survived; the break is in the composite above it.

"Engineers said this could mean that fatigue had weakened the polymer above the brackets, but it could also mean that catastrophic stress had simply snapped the fin. It seemed to rule out festering corrosion at the composite-aluminum joint as the cause of
the casualty.

"Airbus has been using composite tail fins for 16 years, said Roland Thevenin, the
company's certification specialist. The basic test procedure begins with the production of "sample" tail fins containing built-in delaminations and cracks.

"The sample is put through the equivalent of 120,000 "cycles" of a takeoff and landing to
test fatigue and damage tolerance. A "static test" measures composite strength at 1.5
times the maximum load the aircraft is expected to endure. Performance is measured at 70 percent humidity, and at minus 83 degrees Fahrenheit and 158 degrees Fahrenheit.

"When this is finished, the tail fin will not need to be tested again. "If we don't have any [subsequent] visible damage, the structure should be okay for life," Thevenin said. "The typical check is a visual inspection, and if there is something visual, you have to look more closely.""

Edited to add the following excerpts from the November 19 New York Times.

"The tail fin tore in places that investigators said were about two and a half inches thick, and consisted of layers a few thousandths of an inch thick. If a problem in this material is found to be the probable
cause of the crash — a determination is months away — it would be the first crash of a commercial airliner attributed to the failure of a piece of composite, crash investigators say.
....

"It turns out, though, that N14053 needed a repair to its composite tail fin even before it was delivered by the factory in Toulouse,
France, on July 12, 1988. The layers of the tail fin had begun to come apart from within one spot, where the tail connects to the
plane's fuselage.

"When the manufacturer delivered the plane, it said no further inspections of the part were required. The good news was that the
defect, called delaminating, was discovered. But the problem showed that composite can suffer from flaws in its fabrication. That
part of the tail that was fixed appeared to be intact when the tail was examined after the crash."

[ 19 November 2001: Message edited by: SaturnV ]
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Old 19th Nov 2001, 08:57
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Belgique, Thanks for the hint on seeing the enlarged NTSB photos. Yep, the forward attachment(s)? show obvious delaminations; the hard part will be distinguishing the fresh delaminations from the pre-existing delamination(s) -- same for the fractures which are typically at the bolt hole.
With metal, pre-accident fracture surfaces usually show some corrosion/oxydisation while the accident fracture surfaces are shiny. So, how do you determine fracture/delamination age in a particular composite?

Some composite and metal helicopter rotors have crack detection systems; for example, the hollow spaces in the rotor may be pressurised and a pressure drop shows damage somewhere.

In critical areas such as attachment lugs we may need damage detectors. For example narrow foil conductors between laminations would break in the case of a fracture and a structure integrity monitoring system could flag the change in resistance. A mechanism to detect delaminations is also needed. Threading it across layers will be quite a trick; the pre-preg would need a fine hole to enable the detector to traverse the laminations at an angle. The detectors would probably have to be placed before autoclaving.
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Old 19th Nov 2001, 14:08
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"..the fin sheared off above the brackets at its base, neither the brackets nor the bolts anchoring them to the plane's aluminum
fuselage appear to have failed. The joint itself survived; the break is in the composite above it."
Comment: Some are, some aren't (see photos below)
"Engineers said this could mean that fatigue had weakened the polymer above the brackets, but it could also mean that catastrophic stress had simply snapped the fin. It seemed to rule out festering corrosion at the composite-aluminum joint as the cause of
the casualty."
Comment: See remarks below "The Curse of Nylon Peel Ply" and the AD (partly reproduced above)
"When this is finished, the tail fin will not need to be tested again. "If we don't have any [subsequent] visible damage, the structure should be okay for life," Thevenin said. "The typical check is a visual inspection, and if there is something visual, you have to look more closely."
Comment: Statement would appear to be a bit cavalier in the light of what's emerging - but that may well have been the problem all along.

Quite a bit of Boeing's composite comes from Australia. We do the B777 vertical stab ("worlds largest composite component") at the old GAF factory and B737 elevators at the old HdeH - both are now Boeing subsidiaries. HdeH also does flaps for C130J.

We also have the Cooperative Research Centre for Aerospace Structures which develops design and m/f of composites. I've consulted for them - explaining the difficulties of certificating composites. Why we need to be so cautious!
Thanks for alerting me to this AD. Was not aware of it. Sadly keeping up to date gets increasingly difficult.

I'm guessing but this looks like the same problem that occurred during A320 structural testing. And yes, if not corrected, fault might precipitate fin/rudder flutter before it caused an overload failure.

I know of no similar problem being disclosed on the A300.

However the peel ply fault has been lurking for several years. McDD's guru on composites John Hart-Smith (now a Boeing employee and incidentally an Australian) has long preached against dangers of peel ply for the very reasons outlined in the AD.

An example paper by John in conjunction with Redmond and Davis of RAAF Amberley presented in March 1996 is titled "The Curse of Nylon Peel Ply". The abstract begins "A case is presented that most peel plies can be relied upon to ensure the creation of a surface on composite laminates on which it is NOT possible to produce a strong durable adhesive bond.........."

Yet until recently peel ply was industry norm. All manufacturers including McDD used it. I do not know to what extent it has been removed from current manufacturing.

If peel ply gets implicated in the A300 accident it will be real can of worms / chooks roosting time.







Anyone unfamiliar with the flutter phenomenon in composites can look here.
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Old 19th Nov 2001, 15:23
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In Belgique's last post, looking at the top picture, left side composite attachment point, isn't that a row of rivets running from left to right just above the retaining bolt?

Does it appear to anyone else that a fracture occurred at that rivet line? Is this possibly the location where the original delamination occurred at the factory on this fin, and was repaired with rivets and a doubler?
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Old 19th Nov 2001, 16:58
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From looking at all the pictures on the NTSB site the tears and fracture lines in the fin structure suggest that the right side let go first, with the fin tearing away to the left and the aft left attachment letting go last. In the two pictures of the left attachments shown above, there are clearly fasteners through the body of the Centre Left Fin Attachment. It appears therefore, that this is the area of repair mentioned. If that is the case, the final failure sequence seems not to have originated at the repaired attachment. Perhaps the repair was indeed strong enough for the design intent.

Much of the discussion so far has included learned opinions of a lack of knowledge about composite structures and their capabilities. If the state of knowledge is exactly as stated, our aircraft constructors and regulatory bodies are grossly negligent in allowing the use of composite materials in primary structure at all.

In fact, as I sarcastically pointed out on an earlier thread, aircraft manufacturers have built combat aircraft for a long time, using composite primary structures that remain intact at "G" loads far above anything experienced in a civil airliner. These highly loaded combat aircraft do not break apart with any regularity, so the whole problem isn't necessarily poor design or a lack of knowledge about the strength and load bearing capabilities of composite structures. More likely the problem lies in manufacturing techniques and process control. Faced with a delaminated attachment on a complete vertical stabiliser, was it wise to carry out a repair without giving due thought to the reason for the delamination? Was any thought given at all? Obvious delamination on one attachment might have thrown suspicion on the whole assembly. Composite structure is, as has been pointed out, notoriously difficlult to inspect. Did the cost of rejecting and scrapping an entire vertical stabliser drive the repair-and-fit decision? We may well wonder. The NTSB have the difficult task of digging into a foreign manufacturer's production records -but "decision trees" for such decisions are not committed to paper so the answer will be hard to find.

Some years ago, Boeing were building a nice new aircraft for our airline. Some of our Technicians on training courses were visiting the assembly building when one of them noticed a repair on the brand new rudder. Upon inquiry the repair was found to be beyond SRM limits for an in-service rudder - on a new assembly being fitted to a brand new aircraft no less! We asked for the rudder to be removed and replaced with a pristine one. I suspect that it was too expensive to scrap the original unit and I don't know what happened to it after we rejected it. I'd wager though, that it is flying today on an aircraft bought by a less fortunate or observant airline.

We must not lose sight of the real question here. Why was this vertical stabiliser repaired and installed? Is it yet another manifestation of the ever downward pressure on costs, in what has developed into just another low cost transport industry?

**********************************
Through difficulties to the cinema
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Old 19th Nov 2001, 18:42
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From AvFlash:
FLIGHT 587 -- WAKE TURBULENCE IDENTIFIED...
There were some new developments in the investigation of American
Airlines Flight 587, which crashed last Monday shortly after taking off
from New York's John F. Kennedy International Airport. Last Thursday,
NTSB officials announced that they had analyzed the data from the jet's
cockpit and flight data recorders, and investigators believe that the
A300-600 encountered two wake turbulence events caused by a Japan
Airlines Boeing 747-400. Even though standard separation was followed,
data showed sideward accelerations of the aircraft along with
considerable aileron and rudder movements.

...VERTICAL STABILIZER PUZZLES NTSB
The aircraft's vertical stabilizer was recovered from the waters of
Jamaica Bay and found to be largely intact with attachment bolts in
place at the main wreckage site. Wholesale failure of the structure
occurred above the attach points. Investigators are now almost certain
that the tail section was the first part to break off the plane, which
also lost both of its engines before impact. The FAA and French
authorities have ordered airlines to check Airbus A300-600s and A310s
for tail and rudder damage, lose fasteners, distorted surfaces and
cracks and corrosion, within a 15-day period.

NOTE: For more detail of recent findings, see AVweb's Newswire at
<http://avweb.com/n/?47a>.
~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~

Note that the fin's failure was ABOVE the attach points.

Composite failurs?
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Old 19th Nov 2001, 18:51
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Belgique,
Help me out here. If what we are looking at in your pictures is the metal lugs which attach the stabiliser on to the fuselage and if, attached to the lugs are the parts of the stab that remained with the aircraft. How come we see a half rounded hole on the bottom of the stabiliser. Is it because there are four lugs used for the job and we are seeing only two?

Update 15.04
I have actually managed to answer my own question. I now understand that there is a row of lugs at each side of the vertical stab. On closer inspection it appears that one of them is badly damaged. could that be a front lug, if so is it possible that the flutter could have caused the failure of that lug with the subsequent events cascading from there.
http://www.ntsb.gov/events/2001/AA587/AA587_11.jpg

[ 19 November 2001: Message edited by: El Grifo ]

[ 19 November 2001: Message edited by: El Grifo ]
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Old 19th Nov 2001, 19:39
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The FAA just put out an AD on these:

A safety warning affecting the tail section of Airbus A-300 jetliners was scheduled to take effect one day after the crash of American Airlines Flight 587, which plunged out of control Monday after its vertical tail apparently ripped off in flight.

The warning, known as an airworthiness directive, points out a serious safety problem that could leave the pilot without rudder control, making it difficult, if not impossible, to guide the airliner.

The American Airlines jet that crashed after takeoff in New York was an Airbus A-300 Model B4 manufactured in 1988. It was among those that the warning affects.

The U.S. Federal Aviation Authority first published the airworthiness directive on Oct. 29. It was scheduled to take effect Tuesday, forcing all A-300 operators to inspect key rudder-system components. http://www.airsafetyonline.com/cgi-b...MLTemplate?tf= tgam/realtime/fullstory_print.html&cf=tgam/realtime/config-neutral&articleDate=20011117&slug=wxcras&date=20011117&archi v e=RTGAM&site=Front
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Old 19th Nov 2001, 23:28
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Likely to prove a red herring. That story in the Saturday Globe and mail by Peter Cheney can be read here
http://av-info.faa.gov/ad/PublishedADs/012202.html - Airbus Model A300 B2 and B4 Series Airplanes (the rudder AD referred to below)
The A300-600 and A310 are advanced developments of the original A300- B2/B4


Even though this 18 Oct 01 rudder AD is applicable to the accident model Airbus, it's unlikely to have had anything to do with AA587's fate. The facts are slowly emerging that the well-known problem that the A320/A319/A321 model's vertical fin was having (disbonding of the skin from the stringers and spars) was later experienced on the A330 (their much newer product) and that is described here in this AD at this link. Even though this AD does not apply to the A300-B4, the symptoms, as a preliminary to the style of accident AA587 had, were much as predicted in the language of that A320 series AD. The suspicion is that disbonding over a sufficient area can allow the early (gust-induced) initiation of a destructive flutter mode in the vertical fin. The disbonding is caused by polyester contamination introduced during the "peel process" in manufacture (which would be in Spain).

"DGAC advises that localized disbonding has been detected on the skin attachments at the stringers and spars of the spar boxes of the vertical stabilizer. During the manufacturing process, pre-cured parts (attachments of the stringers, spars, and ribs) are installed on the skin panel before the final curing process. A peel ply is used to protect the contact surfaces of the attachment angles of the skin panels of the vertical stabilizer until the pre-cured parts are ready for installation. Investigation revealed that, after the peel ply was removed from
the attachment angles, a residue of polymer finish contaminated the contact surfaces of some pre-cured parts. This contamination
reduced the adhesive strength of the bond and, in some cases, caused debonding (disbonding) of the skin attachments. This condition, if not detected and corrected, could result in reduced structural integrity of the spar boxes of the vertical stabilizer".



This particular rudder problem might lead to a directional controllability issue only if both rudder spring boxes were to seize up, but losing rudder control is nothing at all like having the vertical stabilizer depart the scene. I'd bet money on this one not having been the cause. The reason for the rudder evolutions on the accident aircraft is unlikely to have been pilot inputs and more likely to have been initially the system trying to counter the yaw and roll caused by the vertical fin rocking in a cross-wise fashion (as it fluttered toward detachment after the initial starboard side's "above attachment point" fracture). Latterly the rudder's recorded gyrations may have been due to pre-fin-departure stresses being placed upon the rudder actuator. i.e. As has been pointed out in this thread, the initial flutter mode failure would appear to have been above the attachment points on the starboard side. Once that permitted the flutter to increase in amplitude (laterally), the port side composite lugs, not being designed to cope with those type bending stresses would understandably have fractured through the composite lugs. It doesn't take much imagination to see that.
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Old 19th Nov 2001, 23:44
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At least two red herrings in the newspapers and picked up here:
  • The two fractures in the left center and rear attachment lugs appear well above the bolts and have been seized upon as material defects. Keep in mind that the critical failure was in the forward attachment lugs -- both beginning at the bolt holes where stress would be concentrated (also the right center and rear attachment lugs may also have had pre-accident fractures in the area surrounding the bolts). Once the other four attachments had failed, the fin was coming off. That the failure occurred well above the attachment point testifies to the probable lack of pre-accident fractures in the vicinity of the bolts and an adequate engineering design in the absence of material defects.
  • Speculation that pilot inputs took off the fin. I put my bets on the yaw damper combining with the flight control system. As the forward attachment(s) loosened, the fin leading edge would flex yawing the a/c and the rudder would act to correct the yaw such that the fin leading edge would then flex back and the rudder would then reverse .... resulting in a flutter couple between the structure and the rudder control.
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Old 20th Nov 2001, 00:13
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There have been numerous references to a statement claiming that the VS fractured above the attachment points. Can somebody please reference the original source of this statement? It seems to have evolved into some authoritative claim covering ALL attachment points, and I would like to verify both the 'authoritative' and the 'ALL' presumptions. Furthermore, does this statement include the claim that NO failures occured below the attachment points?

As noted on other posts, a forward bottom attachment appears to have failed. If true, this would give a new perspective on the chain of events.
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Old 20th Nov 2001, 02:44
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The best authority are the photos posted by the NTSB and noted previously in PPRUNE: http://www.ntsb.gov/events/2001/AA587/tailcomp.htm

Look for the three "forward attachment point" photos. The left and right forward lugs are shown along with the forward attachment brackets on the tailcone.

Look at all the photos.
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Old 20th Nov 2001, 03:15
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Stinger?

Sorry to cross post this, but I thought I was posting this here when I initially posted it in the "Newspapers" thread.

OK Don't groan, and this is written by what can only charitably called a lay person.

When we were not at war and with no precedent like the WTC, TWA 800 was firmly believed by most to be a terrorist attack, even to the point that the FBI shouldered the NSTB out of the way for the first few weeks.

Now it is being played in exactly the opposite fashion, though we have precedent and the obvious presence of those with a will to do such a terrible deed.

Also, no one has explained in detail to the public why a hand held SAM type of weapon was not deployed in AA 587. Therefore many are suspecting some sort of cover-up. The lack of any clear definition of the cause of the crash or strongly held theories is also causing concern.

Seeking to dispel such thoughts, one hits the www to find out about Stinger hits on aircraft and is even more confused to find out that China Lake reports that Stinger hits usually do not result in a combustible explosion beyond that of a sharp bright white flash, and drop the target through causing break-up as key parts of the craft shear off which initiates a sharp sudden vertical fall. Of course this does sound like flight AA 587, with the sudden thrust/lurch to the right , the shearing of the tail, and then the dropping away of the engines prior to a large G-force vertical descent.

Could those on this board, especially the many with military flight training behind them, eliminate a SAM as a possible cause of AA 587?

While this may seem silly to request to those who already have such knowledge which seems obvious from their vantage point, it should be reminded that many of your passengers do not share your certainty. It would help if they started to share such certainty as a industrial/parts failure is something that can be understood while a Stinger strike would be a serious problem.

If you feel there is a possibility of a Stinger or similar weaponry strike, why do you think so?

If you can explain with certainty that it is impossible that a Stinger can be the cause, I would love to hear it so we can drive on.

Many thanks.
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