Interesting Airbus / Air France read
06-16-2009, 05:38 PM
#41
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Originally Posted by alvrb211
If you are familiar with the 737 rudder, you'll know that the design is very unusual. Secondly, it was proven to have been the cause of multiple fatal crashes. I'd say it was unusually poor! Even after the modification, it's a very unusual system which lacks redundancy!
06-16-2009, 05:45 PM
#42
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Originally Posted by LivingInMEM
Given how you feel about the 737 rudder design, how do you feel about the use of composite materials in structural areas?
Composites on the 787 will account for 50 percent of the aircraft's structural weight. Aluminum, by contrast, will comprise only 12 percent of the aircraft. In fact, titanium will make up a greater percentage than aluminum, at 15 percent. Steel will comprise another 10 percent and other metals, the remaining 5 percent.
It appears that if there was a lack of confidence in composites, Boeing would not be making an airplane composed of almost 50% composites.
06-17-2009, 08:16 AM
#43
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Originally Posted by III Corps
The original 737 had a dual rudder PCU and yaw damper like the 727 but after tests showed it could be dispatched without the yaw damper, why have two? Thus.. only one for the 737.
No wonder every manufacturer avoids this design.
Al
06-17-2009, 08:55 AM
#45
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Originally Posted by III Corps
It appears that if there was a lack of confidence in composites, Boeing would not be making an airplane composed of almost 50% composites.
Most of the metals that make up the remaining 50 percent are found in very large forgings and castings at the highly loaded joints between composite structures and where the undercarriage is attached.
Although Boeing is treating the frame design as confidential, the one photograph of the frame/fuselage assembly that has been released shows a separate "foot" between the longitudinal hat stringers, to better distribute the shear loads.
The days of the DC-8 and 747 are behind us. 'Estimate the load, then build it twice as strong' is no longer the philosophy.
Last edited by Sniper; 06-17-2009 at 09:00 AM. Reason: added 'saving grace' comment to the end
06-17-2009, 09:03 AM
#46
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Originally Posted by ebl14
Says the guy who will only talk about the 737 in the Airbus A330 crash thread.
wow; just, wow.
wow; just, wow.
Wow huh?
Amid all the rudder talk, someone else raised the B737 problems.
Don't you think it's just a little early to pass judgement in the case of the AF Airbus crash?
Boeing builds a world class product as does Airbus. I'm not bashing either one. In fact, I'm looking forward to seeing Boeing and Rolls Royce launch the B787 in a matter of days.
Al
06-17-2009, 10:43 AM
#47
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Originally Posted by rickair7777
The history of vert stab/rudder failures is on the A300/310 series, not the A320. Different airplanes, although I'm sure some design philosophy bleed over.
In an extreme turbulence situation, the vert stab is probably the most likely thing to fail first on any aircraft. It has a large lever arm.
Also the wings/ailerons and horiz stab/elevators are designed to take large loads in normal flight regime.
The vertical structure normally only needs to handle large control deflections in a low-speed, engine out scenario. It is designed with a lot of area to generate needed counter-yaw at low airspeed.
But there is certainly suspicion that airbus takes a minimalist approach to vert stab strength, as compared to other manufacturers.
In an extreme turbulence situation, the vert stab is probably the most likely thing to fail first on any aircraft. It has a large lever arm.
Also the wings/ailerons and horiz stab/elevators are designed to take large loads in normal flight regime.
The vertical structure normally only needs to handle large control deflections in a low-speed, engine out scenario. It is designed with a lot of area to generate needed counter-yaw at low airspeed.
But there is certainly suspicion that airbus takes a minimalist approach to vert stab strength, as compared to other manufacturers.
06-17-2009, 02:58 PM
#48
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KC-10: Rudder limiters can be defeated - erroneous airspeed indications can be just one prime cause of this.
Vert stabs are certified to full deflection - not rudder reversal as was suspected in the American Airlines accident. Of course, the maximum rudder travel attained on the American Airlines accident was 11 to 12 degrees - hardly a number that I would suspect should result in a structural failure regardless of the airspeed and especially at 251 knots. While this has always been the certification standard, it seems as if the actual strength of the structure has traditionally been more. I would like to see the results from actual static testing on actual complete fuselage structures as to what ultimate load (due to full rudder reversal and side - yaw - loading) catastrophic failure occurs to compare traditional metal structures to the new composite structures. Anecdotally, it seems that we have had more vert stab failures with modern fuselages than with older ones.
There are several reasons that may be. In the past, due to uncertainties in materials science, larger safety factors were built into the computations to account for unknowns (i.e they still designed for 50% safety factors but then added an additional percentage - 20% or so - for materials properties variables). They originally did the same for composites, but have reduced the practice as the science has improved. That is also why composite materials have begun to show up in structural components where they have not before. As the materials science is still new, the engineers can convince managers and manufacturers that they need to spend the extra money on the extra material, but eventually the science gets good enough to the point that the argument for the unknowns starts to fade.
Another reason could be that structures are designed for deflection as well as strength. You simultaneously calculate the required structural member size to handle a given ultimate load prior to failure as well as keep the deflection below a certain maximum value. With certain materials, the minimum dimensions of the member are determined by the maximum allowable deflection and can therefore withstand a higher ultimate load. In this situation, even though the structure is certified to handle that same certain load before failure - it will actually be able to handle a greater load before actual failure (even though the same original design load will exceed that deflection value and may result in permanent deformation, the structure will likely not fail). With composite structures becoming more stiff compared to metal based on the engineering of the materials, the deflection concern is minimized and the ultimate dimensions of the structural members can be reduced.
Personally, I think the manufacturers are advancing the use of composite materials in construction faster than the maintenance technology can keep up. I am still not comfortable with the success rate of current NDI procedures on composite structures. Airlines are still finding delamination and other damage on teardowns of composite components that have recently previously passed visual and NDI inspection. The strength of the composite structure is severely affected by such damage. The effect of long term contact of hydraulic fluids on the strength of the composite material is also virtually incalculable. Between just these two factors, the integrity of the aircraft is being bet on the capabilities of maintenance to mitigate these concerns over the life of the airframe. (Note: Bridges are also designed with particular life spans based on an expected level of preventative maintenance, and we all know what happens when the gov't or owner elects to scale back this maintenance due to economics)
The point is that there is a lot more to aircraft integrity than certification standards and ease and cost of manufacture (the most prominent factor mentioned in that CompositesWorld article posted by III Corps). I will have to do some research to find it, but I am especially curious when Boeing engineers quit due to the extensive use of composites in the 787 structure as happened a year or 3 ago. While the CompositesWorld article was not meant to be an engineering analysis, I would be interested in an objective study that was.
Vert stabs are certified to full deflection - not rudder reversal as was suspected in the American Airlines accident. Of course, the maximum rudder travel attained on the American Airlines accident was 11 to 12 degrees - hardly a number that I would suspect should result in a structural failure regardless of the airspeed and especially at 251 knots. While this has always been the certification standard, it seems as if the actual strength of the structure has traditionally been more. I would like to see the results from actual static testing on actual complete fuselage structures as to what ultimate load (due to full rudder reversal and side - yaw - loading) catastrophic failure occurs to compare traditional metal structures to the new composite structures. Anecdotally, it seems that we have had more vert stab failures with modern fuselages than with older ones.
There are several reasons that may be. In the past, due to uncertainties in materials science, larger safety factors were built into the computations to account for unknowns (i.e they still designed for 50% safety factors but then added an additional percentage - 20% or so - for materials properties variables). They originally did the same for composites, but have reduced the practice as the science has improved. That is also why composite materials have begun to show up in structural components where they have not before. As the materials science is still new, the engineers can convince managers and manufacturers that they need to spend the extra money on the extra material, but eventually the science gets good enough to the point that the argument for the unknowns starts to fade.
Another reason could be that structures are designed for deflection as well as strength. You simultaneously calculate the required structural member size to handle a given ultimate load prior to failure as well as keep the deflection below a certain maximum value. With certain materials, the minimum dimensions of the member are determined by the maximum allowable deflection and can therefore withstand a higher ultimate load. In this situation, even though the structure is certified to handle that same certain load before failure - it will actually be able to handle a greater load before actual failure (even though the same original design load will exceed that deflection value and may result in permanent deformation, the structure will likely not fail). With composite structures becoming more stiff compared to metal based on the engineering of the materials, the deflection concern is minimized and the ultimate dimensions of the structural members can be reduced.
Personally, I think the manufacturers are advancing the use of composite materials in construction faster than the maintenance technology can keep up. I am still not comfortable with the success rate of current NDI procedures on composite structures. Airlines are still finding delamination and other damage on teardowns of composite components that have recently previously passed visual and NDI inspection. The strength of the composite structure is severely affected by such damage. The effect of long term contact of hydraulic fluids on the strength of the composite material is also virtually incalculable. Between just these two factors, the integrity of the aircraft is being bet on the capabilities of maintenance to mitigate these concerns over the life of the airframe. (Note: Bridges are also designed with particular life spans based on an expected level of preventative maintenance, and we all know what happens when the gov't or owner elects to scale back this maintenance due to economics)
The point is that there is a lot more to aircraft integrity than certification standards and ease and cost of manufacture (the most prominent factor mentioned in that CompositesWorld article posted by III Corps). I will have to do some research to find it, but I am especially curious when Boeing engineers quit due to the extensive use of composites in the 787 structure as happened a year or 3 ago. While the CompositesWorld article was not meant to be an engineering analysis, I would be interested in an objective study that was.
Last edited by LivingInMEM; 06-17-2009 at 06:02 PM.
06-17-2009, 03:24 PM
#49
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Originally Posted by Sniper
Neither Boeing, nor Airbus, has figured out a way to use composites to resist the loads at the attachment points yet (due to the way carbon fibers work). .
Maybe they should ask Lockheed (Gen Dyn) how they attached the F-16 verticals. I don't remember the -16s shedding them.
06-17-2009, 03:25 PM
#50
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Originally Posted by KC10 FATboy
Not sure if I disagree with you, but, the vertical stab of an aircraft must withstand a full rudder deflection at any airspeed. If this is not the case, then an rudder deflection limiter is installed -- something the Airbus had.
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