New Clues deepen AA587 Crash Mystery
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During a thirty year airline career from 748s to 747s I cannot recall any dramatic wake turbulence encounter after t/off - sure, a few bumps etc. but that is all.
Has anybody had a severe wake turbulence encounter after t/off in a large a/c?
Has anybody had a severe wake turbulence encounter after t/off in a large a/c?
Keeping Danny in Sandwiches
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I’ve had 2 “interesting” encounters with wake turbulence in 737’s – both approaching in the terminal area. One encounter was along the track of the vortices and caused about a 30 degree bank with a significant amount of opposite aileron along with some rudder input. The other was 90 degrees to the vortices – and caused 2 very abrupt bumps in otherwise calm air. Neither of these prompted a change in draws afterwards but they definitely caught my attention. Of course the 737 doesn’t have the mass of an Airbus 300 but given other structural problems I could see how the event could start the ball rolling.
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A number of people appear to be beginning to question the design choices of the Airbus engineers, apparently on the basis of looking at a few small pictures on the internet.
For the same reason that we would hesitate to start shouting "pilot error" so soon after an accident, let's not start a witch hunt for the engineers until we know a lot more of the facts.
It may turn out that it is down to a design flaw, or something to do with the little known effects of temperature/pressure cycles on composite materials over long periods. One thing is fairly certain: I doubt they would have made such an obvious error as to have made the attachment lugs too small or weak.
In the meantime, a little bit less of the hyperbole please - you don't like it when people do it to you!
For the same reason that we would hesitate to start shouting "pilot error" so soon after an accident, let's not start a witch hunt for the engineers until we know a lot more of the facts.
It may turn out that it is down to a design flaw, or something to do with the little known effects of temperature/pressure cycles on composite materials over long periods. One thing is fairly certain: I doubt they would have made such an obvious error as to have made the attachment lugs too small or weak.
In the meantime, a little bit less of the hyperbole please - you don't like it when people do it to you!
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Not being an engineer myself, though finding this a very interesting subject, I had a chat with a mate of mine the other night over a bottle of wine or three. This guy is responsible for the final approval of the turbine disks in lots of the planes you guys fly, so knows his metallurgy, he doesn't claim to be an expert on composites, but this is what he emailed me today.... What do you say to this Belgique?
"I agree structural composites are a bad idea because repairs are difficult and the lack of toughness of composite materials. I have never seen a repair scheme for a structural composite, because it is not wise and there are few structural composite parts out there. Typically, to manufacture a composite there is the fabric, honeycomb, resin and tooling. The process requires vacuum bagging and a moderate temperature cure cycle with ventilation and plumbing to maintain the vacuum as the bagged composite assembly outgasses - not possible for an in-situ repair. Beyond that composites lack toughness. They have strength to weight ratios greater then metal, but they do not yield. Composites support a load, with little elongation, until failure, which is usually catastrophic - I am not sure what their response to fatigue loading is or their crack propagation characteristics. Because metals are loaded to a fraction of their 0.2% yield strength, they are somewhat easier to design with.
I suspect the tail was composite as a last resort to save weight. The inspection practice for the tail, probably ultrasonic, was insufficient to pickup material degradation. Ultrasonic inspection will pickup voids, but if you have delamination without a void, the sound will pass through the assembly without an echo, and the part will pass the inspection. Other inspection techniques, eddy current or fluorescent penetrant inspection, also require a void of some sort, near surface crack tip or inclusion, or porosity open to the surface.
I believe Airbus have incorporated many technologies to save weight and cost including structural composites and structural casting. But for this there is some risk. Most likely Airbus will have to reevaluate the risk associated with its business model as a result of the accident."
"I agree structural composites are a bad idea because repairs are difficult and the lack of toughness of composite materials. I have never seen a repair scheme for a structural composite, because it is not wise and there are few structural composite parts out there. Typically, to manufacture a composite there is the fabric, honeycomb, resin and tooling. The process requires vacuum bagging and a moderate temperature cure cycle with ventilation and plumbing to maintain the vacuum as the bagged composite assembly outgasses - not possible for an in-situ repair. Beyond that composites lack toughness. They have strength to weight ratios greater then metal, but they do not yield. Composites support a load, with little elongation, until failure, which is usually catastrophic - I am not sure what their response to fatigue loading is or their crack propagation characteristics. Because metals are loaded to a fraction of their 0.2% yield strength, they are somewhat easier to design with.
I suspect the tail was composite as a last resort to save weight. The inspection practice for the tail, probably ultrasonic, was insufficient to pickup material degradation. Ultrasonic inspection will pickup voids, but if you have delamination without a void, the sound will pass through the assembly without an echo, and the part will pass the inspection. Other inspection techniques, eddy current or fluorescent penetrant inspection, also require a void of some sort, near surface crack tip or inclusion, or porosity open to the surface.
I believe Airbus have incorporated many technologies to save weight and cost including structural composites and structural casting. But for this there is some risk. Most likely Airbus will have to reevaluate the risk associated with its business model as a result of the accident."
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As a Metallurgist who became a pilot I can concur with all above.
The bottom line is that although composites have been around for many years they are still relatively new. We simply do not have enough data on the "real" ageing properties of composites. I put that in quote marks because all age data so far is obtained by methods used to prematurely age a material (increasing operating temperature in aluminium being an example). I have a bad feeling that we may now be getting the real data. I hope to god I am wrong and this can be traced to a manufacturing flaw.
16 years ago a professor in this field stated that his opinion was that a composite would lose 50% of its structural integrity over 10 years. Just an opinion but a sobering one.
Another point was that a composite could be stressed repeatedly to 60% of it ultimate tensile strength before suffering fatigue cracking, a metal is typically 30-40% of UTS. In a composite panel a hole(eg a bullet hole) reduces the strengh to 60-70%, in a metal panel this reduces to 90%, a point to ponder in your carbon fibre winged AV8B/Harrier.
My personal feeling is that we are about to see a lot of ageing research carried out on genuinely aged composites.
[ 16 November 2001: Message edited by: whizzjet ]
The bottom line is that although composites have been around for many years they are still relatively new. We simply do not have enough data on the "real" ageing properties of composites. I put that in quote marks because all age data so far is obtained by methods used to prematurely age a material (increasing operating temperature in aluminium being an example). I have a bad feeling that we may now be getting the real data. I hope to god I am wrong and this can be traced to a manufacturing flaw.
16 years ago a professor in this field stated that his opinion was that a composite would lose 50% of its structural integrity over 10 years. Just an opinion but a sobering one.
Another point was that a composite could be stressed repeatedly to 60% of it ultimate tensile strength before suffering fatigue cracking, a metal is typically 30-40% of UTS. In a composite panel a hole(eg a bullet hole) reduces the strengh to 60-70%, in a metal panel this reduces to 90%, a point to ponder in your carbon fibre winged AV8B/Harrier.
My personal feeling is that we are about to see a lot of ageing research carried out on genuinely aged composites.
[ 16 November 2001: Message edited by: whizzjet ]
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Slim-Slag
Couldn't have put it better myself -so I won't. There is a lot more one could say but there's not much point posting it all here. I've written about eight pages on the subject in the last four hours. Data is plentiful, knowledge is not. But enough has been said that I think everyone has the general idea. The AD, if you read it, gives adequate insight into the paucity of inspection and test technologies. It requires you to gaze peerlessly at the areas in question, utilising a strong light if necessary. That sounds like wishful thinking where composite's are concerned but adequately demonstrates that it's an area where the FAA's brightest need more than just a strong light and a magnifying glass. They too will need to call in some fractography experts and buy a machine or two. An extract from a letter to Farley Esq below:
".. (glass-fibre sets exothermically, while carbon-fibre is an endothermic process). When the aviation industry got into the thermo-plasticand thermoset game, they were very cautious inasmuch as it was doors, hatches, fillets and fairings only - for quite a while. They gained experience in laying up composites and the black art of setting the angle of weave of each layer of the cloth matrix became a science. There was a great selection of resins. However when they then crossed the Rubicon into structures they only had that sort of passive panel experience and designing for load-bearing was a suck it and see over-design/over-build process (with lotsa testing to destruction as a confidence builder). Those tests were with aged, as in cured, (but not aging) resins and composites. I owned a Messerschmidt Boelkow Bloehm Phoebus C (VH-GYA) designed by the famed Dr Eppler. It had a mighty balsa substrate with glass overlay and weighed a ton. That airplane came out in 1965 and is still going strong at Cunderdin WA. I reglassed it myself after a number of bingles (never mine) and I was always happier that it was a true hybrid composite. The wing fittings were embedded metal and the all-flying tail's likewise. I could easily live with that but I freely admit that I have great pause for thought when I consider metal brackets hugging composite lugs in the manner shown here and the piccies below. Methinks they have since maybe gotten a little overconfident and that AA587 may just be the proof of that."
I don't endorse the whimsical notion that you'll find now on the front page of aero-news: http://www.aero-news.net/
They are suggesting that there's a mystically strange empennage destructive flutter mode that's peculiar to the A300 and excitated by wake turbulence. I think that the answer's much simpler (as per what has been said already and in particular in the last two posts (by slimslag and whizzjet).
Couldn't have put it better myself -so I won't. There is a lot more one could say but there's not much point posting it all here. I've written about eight pages on the subject in the last four hours. Data is plentiful, knowledge is not. But enough has been said that I think everyone has the general idea. The AD, if you read it, gives adequate insight into the paucity of inspection and test technologies. It requires you to gaze peerlessly at the areas in question, utilising a strong light if necessary. That sounds like wishful thinking where composite's are concerned but adequately demonstrates that it's an area where the FAA's brightest need more than just a strong light and a magnifying glass. They too will need to call in some fractography experts and buy a machine or two. An extract from a letter to Farley Esq below:
".. (glass-fibre sets exothermically, while carbon-fibre is an endothermic process). When the aviation industry got into the thermo-plasticand thermoset game, they were very cautious inasmuch as it was doors, hatches, fillets and fairings only - for quite a while. They gained experience in laying up composites and the black art of setting the angle of weave of each layer of the cloth matrix became a science. There was a great selection of resins. However when they then crossed the Rubicon into structures they only had that sort of passive panel experience and designing for load-bearing was a suck it and see over-design/over-build process (with lotsa testing to destruction as a confidence builder). Those tests were with aged, as in cured, (but not aging) resins and composites. I owned a Messerschmidt Boelkow Bloehm Phoebus C (VH-GYA) designed by the famed Dr Eppler. It had a mighty balsa substrate with glass overlay and weighed a ton. That airplane came out in 1965 and is still going strong at Cunderdin WA. I reglassed it myself after a number of bingles (never mine) and I was always happier that it was a true hybrid composite. The wing fittings were embedded metal and the all-flying tail's likewise. I could easily live with that but I freely admit that I have great pause for thought when I consider metal brackets hugging composite lugs in the manner shown here and the piccies below. Methinks they have since maybe gotten a little overconfident and that AA587 may just be the proof of that."
I don't endorse the whimsical notion that you'll find now on the front page of aero-news: http://www.aero-news.net/
They are suggesting that there's a mystically strange empennage destructive flutter mode that's peculiar to the A300 and excitated by wake turbulence. I think that the answer's much simpler (as per what has been said already and in particular in the last two posts (by slimslag and whizzjet).
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Whilst not qualified to comment on the pro's and con's of the use of composites in aircraft primary structures, I think it is worth pointing out to the anti-airbus brigade that the Boeing 777 also has CFRP primary structure. The vertical stab. primary torsion boxes are CFRP as are the ones in the horizontal stab. If this does prove to be a composite failure, then it is of great concern not only to Airbus but the whole aviation industry.
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Lame - Thanks for your info - I didn't Know that! Gooday to you!
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Belgique : glass sets exothermically wheras carbon endothermically ?. It is the resin that cures and in most cases a simlar epoxy resin is used for Carbon fibre and glass fibre. As far as I know all epoxies require heat to cure and will generate exothermic heat as the cross linking process goes on ( in thick laminates care has to be taken with the heat up rate to avoid run away exothermic reaction ). Can you expand a bit on the statement.
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Some years ago I took off behind an MD-11, I was flying a fully loaded DC-8 freighter. Separation was not more than a minute. At about 800 feet, the aircraft began to roll starboard, full control wheel throw in the opposite direction didn’t stop it, merely slowed it down somewhat. At 40 degrees of bank, we exited the vortex and recovered.
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I reckon Belgique is on the right track, there resides in the fair place in which he (or at least his aircraft anyway) resides being the home of a fair bit if research in this area.
Having said that and looking at the attach diagram, what appears to be missing are multiple load paths to the main structure, ie the absence of a "fail safe" mechanism.
OK OK there are three brackets but the descriptions of the failure mode I have heard so far seem to suggest that the loss of any bracket but the middle one would tend to accelerate the separation. This seems to me to be a bit dodgy.
There does not appear to be (and I am not an engineer) an alternative load path from the VS to the main structure beyond those three brackets??
Belt and braces, sure, but I thought that was what the certification process had evolved towards and is this still a requirement.
This is not a criticism of the product or the magnificent technology, but when the Governments, therefore certifying authority and manufacturers are the same people and national pride is involved the pressure towards pushing the boundaries has to be higher.
IMHO the entire raison d'etre of the Airbus philosophy is towards maximising modern lightweight composite technology which awards better operating economics against the higher order of technical difficulties in manufacture and repair. A more conservative approach might produce weight penalties that nullify those benefits.
This must inevitably lead to a sort of technical hubris.
I noticed some one talking about composite primary structure in the B777, I would be interested to know if they also provide an alternative load path around it and/or what are the consequences of it failing?
And no I don't work for Boeing.
allianceair
No I have not found the photo yet but would appreciate some help. I did ask if anyone else had seen it.
[ 17 November 2001: Message edited by: gaunty ]
Having said that and looking at the attach diagram, what appears to be missing are multiple load paths to the main structure, ie the absence of a "fail safe" mechanism.
OK OK there are three brackets but the descriptions of the failure mode I have heard so far seem to suggest that the loss of any bracket but the middle one would tend to accelerate the separation. This seems to me to be a bit dodgy.
There does not appear to be (and I am not an engineer) an alternative load path from the VS to the main structure beyond those three brackets??
Belt and braces, sure, but I thought that was what the certification process had evolved towards and is this still a requirement.
This is not a criticism of the product or the magnificent technology, but when the Governments, therefore certifying authority and manufacturers are the same people and national pride is involved the pressure towards pushing the boundaries has to be higher.
IMHO the entire raison d'etre of the Airbus philosophy is towards maximising modern lightweight composite technology which awards better operating economics against the higher order of technical difficulties in manufacture and repair. A more conservative approach might produce weight penalties that nullify those benefits.
This must inevitably lead to a sort of technical hubris.
I noticed some one talking about composite primary structure in the B777, I would be interested to know if they also provide an alternative load path around it and/or what are the consequences of it failing?
And no I don't work for Boeing.
allianceair
No I have not found the photo yet but would appreciate some help. I did ask if anyone else had seen it.
[ 17 November 2001: Message edited by: gaunty ]
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Since the aeronautical engineering world allegedly knows b*gger-all about structural composites, perhaps its time for all those Harrier/AV8 and certain other military pilots to ease up on the aerobatics. I mean, pulling +9G/-3Gs they're more likely to have the wings fall off rather than getting them shot off by anti-aircraft fire.
Or aren't they?
Thank goodness aeroplanes aren't designed in bars over a few pints...
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Or aren't they?
Thank goodness aeroplanes aren't designed in bars over a few pints...
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Not an engineer, mechanic, pilot, anything... BUT, I'm tellin ya...too much attention to composite removes the visual clues. Aft left attachment seems to indicate a forward right to aft left removal, leading one to look at the forward right attachments. It's a fuzzy picture, but the BOTTOM forward right bracket seems to be bent up and back. Leads me to think it was that bracket, a metal bracket, which failed and caused the 'rocking' so well described by Belgique - leading to failure of top composite brackets.
http://www.ntsb.gov/events/2001/AA587/AA587_11.jpg
Going crazy, thinking I'm crazy. Somebody (vastly more knowledgable then I) please give some technical advise on why the attached picture doesn't show what I think it shows.
...not trying to defend Airbus or composites...
http://www.ntsb.gov/events/2001/AA587/AA587_11.jpg
Going crazy, thinking I'm crazy. Somebody (vastly more knowledgable then I) please give some technical advise on why the attached picture doesn't show what I think it shows.
...not trying to defend Airbus or composites...
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Unfortunately, due to lax statements by NTSB reps. and the usual ignorant press reporters the pilots are already being blamed for this accident in the US press. Along the lines of "actions of the pilots caused the rudder to fail".
I would welcome opinions on my thoughts. Following a couple of loss of control events our company started, quite rightly, to train recovers from upsets. At some point the use of rudder became a PRIMARY recovery aid to be used in all circumstances where roll enhancement is needed. I have always felt that the rudder should be used with extreme caution because of the danger of overstress. Unfortunately, because we have never had a rudder break off during simulator training we are, I believe, encouraging pilots to take actions which could lead to rudder overstress during flight. I sat in on simulator training recently where FULL rudder was applied at 300 + knots!
What do other airlines teach?
I would welcome opinions on my thoughts. Following a couple of loss of control events our company started, quite rightly, to train recovers from upsets. At some point the use of rudder became a PRIMARY recovery aid to be used in all circumstances where roll enhancement is needed. I have always felt that the rudder should be used with extreme caution because of the danger of overstress. Unfortunately, because we have never had a rudder break off during simulator training we are, I believe, encouraging pilots to take actions which could lead to rudder overstress during flight. I sat in on simulator training recently where FULL rudder was applied at 300 + knots!
What do other airlines teach?
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I just don't get it. The FAA Airworthiness Directive seems terribly inadequate. All the posts on this and other forums on composite structures tell of the difficulty in adequately (non-destructive) testing of composite material. How is a cursive visual inspection (literaly "lights" and "mirrors") going to determine the structural integrity of the vertical stab?
I would have thought that perhaps the newer aircraft in the fleet could initially be subject to less scrutiny, but the older higher time/cycle aircraft be inspected more thouroughly.
Comments from some engineering types?
I would have thought that perhaps the newer aircraft in the fleet could initially be subject to less scrutiny, but the older higher time/cycle aircraft be inspected more thouroughly.
Comments from some engineering types?