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Cfrp Panel Fuselages Superior (?)  
User currently offlineRheinwaldner From Switzerland, joined Jan 2008, 2241 posts, RR: 5
Posted (6 years 8 months 1 week 3 days ago) and read 24571 times:

In the past endless discussion were on A-Net about the composite fuselage construction:

RE: Airbus Sticks With Panels For The A350 (by WingedMigrator Jun 9 2007 in Civil Aviation)
Airbus A350 Cfrp Panel Construction (by WingedMigrator Jun 22 2007 in Tech Ops)

A350XWB - Back To The Drawing Board (again)? (by N1786b May 29 2007 in Civil Aviation)
A350XWB - Back To The Drawing Board (again) Pt 2 (by ANCFlyer May 31 2007 in Civil Aviation)

Confirmed: Composite Frame On A350 XWB (by Keesje Sep 20 2007 in Civil Aviation)


I closely watched these threads and found following consent:
The Barrel approach (B787) is generally considered superior versus the Panel approach (A350).
Among the many advantages following are mentioned often:


  • Cheaper and fully automated production
  • High integration from the beginning
  • Lighter (because of the following two points)
  • Less fasteners
  • Fewer joints


In this thread I want to discuss why Airbus is on the Panel track. I will support this by substantial sources and ideas. Among the most heard reasons why Airbus plans to use Panels we often heard so far:

  • Airbus has not the composite know how
  • Airbus can not use Barrels because of intellectual property issues
  • Airbus needs to go the fast approach therefore has not the time to make it 'right'
  • Airbus has not the logistics to transport Barrels


These all imply that Airbus needs to choose a technologically inferior solution because of surrounding requirements. I don't want to add much comment on these but want to raise the question: could there also be technological advantages that we just don't know?
I just mention in short the possible advantages that were listed so far:

  • Easier logistics when transporting panels instead of barrels
  • Longer panels than barrels possible


When seeking for other advantages quite often there were gentle hints by WingedMigrator but usualy he only earned ingorance or even opposition ( Airbus A350 Cfrp Panel Construction (by WingedMigrator Jun 22 2007 in Tech Ops) ). First I want to share what I gathered around of what he suggested. Let's develop a line of reasoning:
 idea 

  • With Panels the outside of the fuselage lay in the Mandrel. Therefore you will get the smother surface on the aerodynamically important side of the wall (or at least better control on the outcome during manufacturing). Kaneporta1 once raised this point.
  • With barrels on the inside you can only vary the thickness to some degree and have limited possibilities to built 3D structures (for the B787 only stringers)
  • Having the outside in the form you are free to control the structure complexity on the inside and built 3D structures.
  • Thus you could have the idea to integrate and co cure various inner construction elements from the beginning. This could be:

    • Stringers
    • Frames
    • Window Frames
    • Door frames
    • Reinforcements that can not be built only by thickness change
    • Crash absorbing construction elements

  • I made a graphic that shows a panel with fully integrated stringers and frames:
    It would come as one piece out of the oven. The number of fasteners is greatly reduced. The splice plate on one side could be made from the panel itself thus eliminating further rows of fasteners.


  • The built process to produce such panels would be fully automated. For final assembly the four panels just need to be connected with long rows of fasteners, and that's it (moderate handwork to get a complete fuselage barrel made out of 4 parts). In contrast with the barrel approach you get little more than the skin and you need to put in a large number of frames (more handwork to get a complete fuselage barrel consisting of many more than 4 parts).
  • On the B787 only the stringers are integrated into the barrels. As a result a fastener orgy lead to pictures as these:
    http://www.airliners.net/uf/view.fil...2901&filename=1179807132iuzSrA.jpg
    http://www.airliners.net/uf/view.fil...2901&filename=1179808008hNFyUy.jpg
    If we compare the required number of fasteners the panel approach IS the winner. See this picture: http://www.vought.com/gallery/locati.../southCarolina/sc_production18.jpg -> All the white little spots are fasteners that could be left away with the co-cured panel approach.
  • The next picture shows a segment of a fuselage with one barrel joint. Red dots are fasteners.
    Picture 1: Barrel approach
    Picture 2: Panel approach light (only frames and stringers co cured with the panel)
    Picture 3: Panel approach max (Stringers, frames, doorframes, window-frames, reinforcements all co cured with the panel)



Conclusion:
In the light of these ideas the picture changes dramatically. From the listed advantages for the barrel approach (easier production, Lighter, Less fasteners, Fewer joints) each one can be disposed. The panel approach leads to higher integrated parts allowing cheaper (more automated) production and final assembly. There are fewer fasteners and joints (if you count the joint between each frame and the barrel too).

Availability:
You see the panel has the clear potential to be technically superior. The question remains whether today composite technology is mature enough to produce such a thing (how complex can structures be built up on the panel inside?). If one day the industry succeeds to realize the "Panel approach max" as outlined above you will for sure see Boeing change to the panel approach!

Now how much of this technology is at hand for Airbus to produce such structures?
So far in this post I summarized what A-Net has brought up so far and what the advantages of panels could be. In the second part of the post I want to share all the sources I have found around this topic. A few google-searches yield a lot of interesting stuff regarding this question. I have found some treasuries of links that shed a lot of light on these questions. A general review of public information available on the internet allows following statements:
- There is a huge amount of know how for composite fuselages available in Europe
- Also crash worthiness is covered well with studies

The first link reveals about a crash worthiness study made in Holland in 2001:
http://nlr.nl/smartsite.dws?id=4366
Very similarly to the Boeing 787 crash-tests a composite fuselage section was crashed to verify the accuracy of computer simulations. The report shows that the predictions were not top but still the tests showed principles to make the results predictable.
Remarkable is following structural concept: "The ring-frame configuration as commonly used in fixed-wing aircraft, is a difficult component for crashworthiness. NASA studies have shown, that the "point-load" applied by the ground leads to immediate fracture at this point, followed by severe bending, and further breakages of the frame higher up [7]. This may result in an early disintegration of the structure, and the bypassing of the dedicated energy-absorbing measures. The recommendations followed from these studies were, to separate the livable volume "on top" from the energy absorbing components "below". Hence, a sufficiently strong, closed ring-frame, meant to survive the impact, should be positioned on top of an expandable energy absorbing structure...."
In summary: Having a separated upper and lower structure in fuselage allows in theory a better design regarding crash-worthiness. And regarding the panel-advantage-discussion: By using panels instead of a barrel this concept can be implemented better.

The second link gives deep insight into European composite fuselage research:
http://www.dlr.de/wf/en/Portaldata/2...t-fuer-einen-cfk-flugzeugrumpf.pdf
Document review:
Page 5 below: Sidenote about new metal technologies: they promise to have the same potential as today composites regarding cost and weight. But, future composites promise to be unbeatable. Therefore second generation composite fuselages will have a clear advantage over what is top today.
Page 6 below: Possible concepts to make composites conductive (hint: improved lightning protection)
Page 8 below: As above a separation of lower and upper fuselage functions.
Page 9 Top: Shell concept with integrated stringers. Design with co cured stringers and frames is achieved.
Page 11+12: Superior crash worthiness of Gondola concept
Page 12 below: Innovative and cheaper manufacturing, reduced number of bolts
Page 13 below: Summary of various European composite fuselage research programs
General: Work seems more concrete for an A320 sized fuselage, even a mixed demonstrator is mentioned on A320 basis. Building an A320 with a composite fuselage is therefore not impossible.

Now what do you think? Will the panels for the A350 already be superior to any barrel approach?

Forecasts
Based on all of this and to stir up the discussion I want to make some forecasts. These are only personal opinions and I will freely admit my wrong guess if the opposite turns to be the truth:
 eyepopping 
Forecast 1: CFRP Barrel fuselages will turn out to be inferior regarding crash worthiness
Forecast 2: CFRP Barrel fuselages will turn out to be more expensive in production
Forecast 3: CFRP Barrel fuselages will turn out to be heavier
Forecast 4: The B787 will be the last airliner using CFRP barrel fuselages
Forecast 5: The NG Narrowbodies from Boeing and Airbus will use CFRP panels
Forecast 6: The first 787 write off will be after an incident smaller than BA038 because of unrepairable fuselage damage (something like this:
View Large View Medium
Click here for bigger photo!

Photo © Jarett Sirko

, such a hole can not be patched)
Forecast 7: After the first comparing incident with an A350 the plane will not be written off
Forecast 8: CFRP Panel fuselages will allow replacement of single panels for repair of heavy damage (how?: Put the plane in a corset to fix the shape while part of the skin is missing, remove the fasteners for the panel, skin with stringers and frames is replaced as one piece).

107 replies: All unread, showing first 25:
 
User currently offlineAutoThrust From Switzerland, joined Jun 2006, 1596 posts, RR: 9
Reply 1, posted (6 years 8 months 1 week 2 days 20 hours ago) and read 24442 times:

Very interesting post, Rheinwaldner. I'm no expert, but a other reason for shells are the need of smaller autoclave's unlike Boeing. You need gigantic autoclaves for Barrels.


Btw Airbus is doing research on a EU project called MAAXIMUS which include advancing simulation-based composite airframe design, improving manufacturing technology, and achieving a reduction of 10% in structural weight and 20% in development lead time, all by 2012.(could be applied on the A320RS)

http://www.aerosme.com/download/Work...ll/docs/WebVersion/08_MAAXIMUS.pdf



“Faliure is not an option.”
User currently offlineHawkerCamm From United Kingdom, joined Jul 2007, 405 posts, RR: 0
Reply 2, posted (6 years 8 months 1 week 2 days 19 hours ago) and read 24351 times:

Excellent post and a very good technically analysis.

The mould lines being on the outside of the fuselage should also led to a very smooth external finish. The B787 barrels results in a fairly rough external finish


User currently offlineArniePie From Belgium, joined Aug 2005, 1265 posts, RR: 1
Reply 3, posted (6 years 8 months 1 week 2 days 19 hours ago) and read 24296 times:



Quoting Rheinwaldner (Thread starter):
Forecast 6: The first 787 write off will be after an incident smaller than BA038 because of unrepairable fuselage damage (something like this:
View Large View Medium
Click here for bigger photo!
Photo © Jarett Sirko

, such a hole can not be patched)

Overall some interesting remarks you make about the different methods of how to make the best Carbon type fuselage.

I don't want to comment on all conclusions you made but point 6 is certainly not true, even major damage as that (like a collision with a service truck or something similar) should be repairable when you use a barrel.
The patch (basically a custom made panel) could add some weight to the airframe but it should be minimal.



[edit post]
User currently offlineTdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 4, posted (6 years 8 months 1 week 2 days 17 hours ago) and read 24193 times:



Quoting Rheinwaldner (Thread starter):
could there also be technological advantages that we just don't know?
I just mention in short the possible advantages that were listed so far:


* Easier logistics when transporting panels instead of barrels
* Longer panels than barrels possible

The logistics are definitely easier, but I don't believe the second point is true. The limit on length of panels or barrels today is the length of the autoclave...a 100' autoclave can only make a 100' panel or barrel. Diameter isn't as much of an issue as you might think since, for pressure reasons, autoclaves are almost always round anyway.

Quoting Rheinwaldner (Thread starter):
With Panels the outside of the fuselage lay in the Mandrel. Therefore you will get the smother surface on the aerodynamically important side of the wall (or at least better control on the outcome during manufacturing).

There is no particular reason you can't do a barrel with a tooling surface on the outside, although I don't believe the 787 is doing so. I don't think this is an inherent panel vs. barrel issue.

Quoting Rheinwaldner (Thread starter):
With barrels on the inside you can only vary the thickness to some degree and have limited possibilities to built 3D structures (for the B787 only stringers)

Provided you're willing to make the mandrel more complex, you can put structures and thickness changes to almost arbitrary degree for panels or barrels.

Quoting Rheinwaldner (Thread starter):
#
# I made a graphic that shows a panel with fully integrated stringers and frames:

There's a slight issue here with your lap joint...tremendous off-center loading (and a heck of a stress riser in the sharp corner). Although I certainly agree you could integrate the lap joint into the panel, it strongly doubt the cross section would look like this.

Quoting Rheinwaldner (Thread starter):
The panel approach leads to higher integrated parts allowing cheaper (more automated) production and final assembly. There are fewer fasteners and joints (if you count the joint between each frame and the barrel too).

Both panel and barrel construction can support higher integration than either Airbus or Boeing are currently considering. I'm don't see why your point above applies to panels and not barrels.

Quoting Rheinwaldner (Thread starter):
Having a separated upper and lower structure in fuselage allows in theory a better design regarding crash-worthiness. And regarding the panel-advantage-discussion: By using panels instead of a barrel this concept can be implemented better.

Again, not sure why. One of the huge advantages of composites, in both panel and barrel form, is local tailoring of structural properties. You can build a barrel with very different upper and lower crash response just as you can build it with panels.

Quoting Rheinwaldner (Thread starter):
Forecast 1: CFRP Barrel fuselages will turn out to be inferior regarding crash worthiness

As a general statement, I believe this is false because of the point raised above.

Quoting Rheinwaldner (Thread starter):
Forecast 2: CFRP Barrel fuselages will turn out to be more expensive in production

Absolutely true on a part basis. For an overall airplane cost basis, I doubt it because of the savings in final assembly joints.

Quoting Rheinwaldner (Thread starter):
Forecast 3: CFRP Barrel fuselages will turn out to be heavier

Highly doubtful given that, no matter how you slice it, panels will have more major joints.

Quoting Rheinwaldner (Thread starter):
Forecast 6: The first 787 write off will be after an incident smaller than BA038 because of unrepairable fuselage damage

Doubtful. Repairability is negligably different between the two technologies because for anything up to a full panel replacement, you're going to be doing the same patch on either one. If you go to full panel replacement (a *major* repair) the equivalent patch on a barrel is of the same complexity.

Quoting Rheinwaldner (Thread starter):
Forecast 8: CFRP Panel fuselages will allow replacement of single panels for repair of heavy damage (how?: Put the plane in a corset to fix the shape while part of the skin is missing, remove the fasteners for the panel, skin with stringers and frames is replaced as one piece).

This is certainly true, and is equivalent to a reskinning job today. However, this is a huge repair and isn't undertaken lightly. A tiny fraction of a percent of all fuselage repairs over the life of an entire fleet are done by skin replacement, so the value amortized over the fleet is relatively small.

Tom.


User currently offlineRedFlyer From United States of America, joined Feb 2005, 4335 posts, RR: 28
Reply 5, posted (6 years 8 months 1 week 2 days 15 hours ago) and read 24137 times:

Very interesting thread.

Wouldn't the Panel Max approach actually increase manufacturing costs? It sure looks "prettier" and cleaner from a fastener standpoint, but I imagine co-curing the stringers, frames, doorframes, and window-frames is going to add a considerable level of complexity to the manufacturing process. The forms (press) are going to have to be very precise within very tight tolerances and all the while accounting for a multitude of complex shapes and angles. More to the point, would it even be possible to cure a panel containing all of those different parts at one time or would the panel have to undergo multiple curings? Wouldn't one also require a unique form or press for each panel? Would that also entail an autoclave for each form or press?

I know, a lot of questions. I'm just an ignorant layman, but that is what popped into my mind when I read throught the OP's post.



My other home is a Piper Cherokee 180C
User currently offlineTdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 6, posted (6 years 8 months 1 week 2 days 14 hours ago) and read 24113 times:



Quoting RedFlyer (Reply 5):
Wouldn't the Panel Max approach actually increase manufacturing costs?

As compared to a regular panel, probably, although it would be offset somewhat by reduced assembly costs.

Quoting RedFlyer (Reply 5):
More to the point, would it even be possible to cure a panel containing all of those different parts at one time or would the panel have to undergo multiple curings?

Unlikely, due to handling problems if nothing else. It would be easier to partially cure the parts individually to give them some structure and allow handling without them distorting, then do the final co-cure to bond it all together.

Quoting RedFlyer (Reply 5):
Wouldn't one also require a unique form or press for each panel?

Unique tooling or reconfigurable tooling. Probably no press involved (that's what the autoclave is for).

Quoting RedFlyer (Reply 5):
Would that also entail an autoclave for each form or press?

Probably not...the autoclave is just a big pressurized oven. Provided it's big enough for your parts and tooling to fit in, you can use it over and over on different parts (or cure multiple parts together, if the cure cycle is the same).

Tom.


User currently offlineRedFlyer From United States of America, joined Feb 2005, 4335 posts, RR: 28
Reply 7, posted (6 years 8 months 1 week 2 days 6 hours ago) and read 24057 times:



Quoting Tdscanuck (Reply 6):
Quoting RedFlyer (Reply 5):
More to the point, would it even be possible to cure a panel containing all of those different parts at one time or would the panel have to undergo multiple curings?

Unlikely, due to handling problems if nothing else. It would be easier to partially cure the parts individually to give them some structure and allow handling without them distorting, then do the final co-cure to bond it all together.

That was my whole point: "partially curing the parts individually...then do the final co-cure" would seem to be multiple curings. All of which adds time and complexity to the process. It certainly doesn't make it more efficient. The individual parts will still have to be "constructed" prior to the final sub-assembly. Rather than construct the support structures on the back-end of the process as Boeing does on its Barrrel method, Airbus would construct the support structures on the front-end of their Panel method.

Quoting Tdscanuck (Reply 6):
Quoting RedFlyer (Reply 5):
Wouldn't one also require a unique form or press for each panel?

Unique tooling or reconfigurable tooling. Probably no press involved (that's what the autoclave is for).

Just curious, but how would the parts be held in place during the autoclave process? I was under the impression an autoclave applies pressure and heat to "cure" the compound by removing all air from the material. How would the separate parts be applied (all at the same time, I presume) under pressure while ensuring they remain static? (That was why I used the term "press".)

Quoting Tdscanuck (Reply 6):
Quoting RedFlyer (Reply 5):
Would that also entail an autoclave for each form or press?

Probably not...the autoclave is just a big pressurized oven. Provided it's big enough for your parts and tooling to fit in, you can use it over and over on different parts (or cure multiple parts together, if the cure cycle is the same).

I could see using a single autoclave for different panel parts. But that would kind of slow things down a little, wouldn't it? I believe we're talking roughly a dozen panels that would comprise the fuselage. That's 12 individual "cooks" that have to take place. Seems it would remove a lot of the efficiency if that has to be done (and we won't even go into defective panels that have to be re-done). Unless, of course, Airbus were to outsource the panels to different vendors who could cure panels at the same time at different locations.

In any event, the OP's original post is eye-opening. I can see some of the advantages to the Panel approach (the Panel Max option). I don't know if it's superior than the Barrel method, but definitely better than I thought it was before.



My other home is a Piper Cherokee 180C
User currently offlineWingedMigrator From United States of America, joined Oct 2005, 2214 posts, RR: 56
Reply 8, posted (6 years 8 months 1 week 2 days 5 hours ago) and read 24052 times:

Thanks Rheinwaldner for the interesting post. Your picture 3 "panel max" is indeed what I had been inquiring about, without success, and you explained it and illustrated it much more clearly.

I don't follow your conclusions on the superiority of the panel approach... in my view there are too many variables and unknowns that prevent the categorical statement that one or the other is better.

But you are correct that the barrel approach, being just one of several possible approaches, is sometimes oversold. Or, more to the point, the panel approach is often dismissed as inefficient, adding what sounds like many tons -- when in fact these bolted joints are a tiny fraction of the mass of the finished airplane. All this is very far from being clear at this point.

Quoting RedFlyer (Reply 5):
I imagine co-curing the stringers, frames, doorframes, and window-frames is going to add a considerable level of complexity to the manufacturing process.

Yes... but Airbus seems to have some experience in this regard. For example, the A380 rear pressure bulkhead (or rather, Bulkhead with a big 'B') is a monolithic CFRP structure that undergoes at least 2 cure cycles, one for the skin and the second for co-curing ribs. Perhaps some of this complexity could be offset by the simpler assembly.

This has a picture of it (see figure 2-17)
http://dspace.lib.cranfield.ac.uk:80...tream/1826/1657/1/Goachet-2006.pdf

(edit: here's another better link describing the manufacturing process of the A380 rear pressure bulkhead in detail... http://www.compositesworld.com/hpc/issues/2003/May/101)

Quoting RedFlyer (Reply 7):
The individual parts will still have to be "constructed" prior to the final sub-assembly.

By machines, perhaps? I'm not sure why this would be a disadvantage.

Quoting RedFlyer (Reply 7):
I believe we're talking roughly a dozen panels that would comprise the fuselage. That's 12 individual "cooks" that have to take place.

You are possibly assuming that only one panel is allowed per autoclave cycle. I see nothing in the panel approach that prevents a rack of several panels from being cured simultaneously.

Quoting RedFlyer (Reply 7):
and we won't even go into defective panels that have to be re-done

That's potentially a problem with barrels as well, since losing a barrel is like losing four panels at once. If the yield were significantly less than 100%, panels would be advantageous, but everything I've heard about the Boeing operation suggests that yield is very close to 100%.

[Edited 2008-01-25 21:09:22]

User currently offlineRedFlyer From United States of America, joined Feb 2005, 4335 posts, RR: 28
Reply 9, posted (6 years 8 months 1 week 2 days 2 hours ago) and read 24027 times:

Thanks for keeping this thread going, WingedMigrator. I find it one of the more interesting so I'm looking for more detailed explanations. On to your specific points...

Quoting WingedMigrator (Reply 8):
Quoting RedFlyer (Reply 5):
I imagine co-curing the stringers, frames, doorframes, and window-frames is going to add a considerable level of complexity to the manufacturing process.

Yes... but Airbus seems to have some experience in this regard.

No doubt Airbus has experience with this and they could do it very efficiently. What I'm curious to know is how this specifically might be a better (read more efficient) process than Boeing's barrel approach. As I said, Boeing is constructing the stringers, frames, door frames, and window frames on the back-end of the construction process. So Airbus does it on the front end - how does that make it more efficient? Parity, I can understand. I'm not sure I grasp the "more" efficient aspect of it.

Quoting WingedMigrator (Reply 8):
Quoting RedFlyer (Reply 7):
The individual parts will still have to be "constructed" prior to the final sub-assembly.

By machines, perhaps? I'm not sure why this would be a disadvantage.

I would think a lot of the sub-components (stringers, frames, doorframes, and window frames) are also machined on the Boeing side. Don't forget: most of the parts and sub-assemblies by both manufacturers are machine (robot) made these days.

Quoting WingedMigrator (Reply 8):
Quoting RedFlyer (Reply 7):
I believe we're talking roughly a dozen panels that would comprise the fuselage. That's 12 individual "cooks" that have to take place.

You are possibly assuming that only one panel is allowed per autoclave cycle. I see nothing in the panel approach that prevents a rack of several panels from being cured simultaneously.

Ah, definitely a big step forward! A large-enough autoclave could have several "shelves" or racks that accommodate multiple panels and that would definitely make it a more efficient process.

Quoting WingedMigrator (Reply 8):
Quoting RedFlyer (Reply 7):
and we won't even go into defective panels that have to be re-done

That's potentially a problem with barrels as well, since losing a barrel is like losing four panels at once. If the yield were significantly less than 100%, panels would be advantageous, but everything I've heard about the Boeing operation suggests that yield is very close to 100%.

Ok, I see your point; however, in manufacturing "defects" are defined by defects per "X" number of parts. If we assume all of the sub-components (stringers, frames, etc.) are a wash between the two methods since they will both require manufacture of those components in similar quantities separately for inclusion in the final subassembly products, then the barrel method actually has LESS parts than the panel method. That's because currently Boeing constructs 7 barrels per fuselage (I believe their goal is to eventually reduce this number to 4 or even 3 barrels). The panel method would require roughly a dozen panels (parts) for a fuselage. If we are to assume 1 defect per 100 parts then that means that every 1 of every 8 airframes will encounter a defective part requiring a re-do. On the barrel method, however, 1 defect per 100 parts translates into 1 defective part for every 14 airframes. (This is sounding an awful lot like that age-old A.net argument about twins vs. quads!) Of course, I'm assuming 1 in 100 defects - it could in fact be a lot better ratio which could render it a relatively moot point. And, of course, people could point out that a mistake on a Boeing barrel part, while less likely, is more costly because it's a bigger part (more materials, more time to construct, etc.).

(BTW, neither of the links you provided worked on my end.)

[Edited 2008-01-26 00:00:56]


My other home is a Piper Cherokee 180C
User currently offlineWingedMigrator From United States of America, joined Oct 2005, 2214 posts, RR: 56
Reply 10, posted (6 years 8 months 1 week 1 day 16 hours ago) and read 23967 times:



Quoting RedFlyer (Reply 9):
What I'm curious to know is how this specifically might be a better (read more efficient) process than Boeing's barrel approach.

I haven't claimed that one was superior to the other. In my view the pros and cons of each method are not sufficiently well-defined to conclude positively one way or the other. Others do conclude... Rheinwaldner is in the 'panel' column, and lots of 787 fans are firmly in the 'barrel' column, but I don't think either side has made a convincing case that a clear advantage exists.

Quoting RedFlyer (Reply 9):
As I said, Boeing is constructing the stringers, frames, door frames, and window frames on the back-end of the construction process.

Boeing's stringers are integral to the barrels. Everything else is, as you say, manually installed inside the barrels with fasteners. Co-curing all that at the front end of the process may further cut down on the amount of touch labor. The panel approach may eliminate as many fastened joints as it adds.

Quoting RedFlyer (Reply 9):
in manufacturing "defects" are defined by defects per "X" number of parts.

One thing to keep in mind is that "X" probably depends on the size of the part. This is true of microchips and computer displays: the larger the part, the lower the yield, because the defects occur at a given rate on a scale much smaller than the whole part. I don't know if this is true of large CFRP parts-- as I said, nothing indicates that Boeing's yield is suffering from the large size of their parts.

Quoting RedFlyer (Reply 9):
neither of the links you provided worked on my end

That's really too bad. This link from 2003 has a fascinating, detailed and illustrated description of the manufacturing steps for the A380 bulkhead. They mention that eventually, reinforcing features could be co-cured in a single autoclave cycle, but they are sticking with two cycles for now because they want to stay within their experience and lower risk.

It's not too difficult to imagine a similar manufacturing process for the A350 fuselage panels, in the "Panel Max" configuration described by Rheinwaldner above.


User currently offline474218 From United States of America, joined Oct 2005, 6340 posts, RR: 9
Reply 11, posted (6 years 8 months 1 week 1 day 14 hours ago) and read 23952 times:



Quoting Rheinwaldner (Thread starter):
I made a graphic that shows a panel with fully integrated stringers and frames:
It would come as one piece out of the oven. The number of fasteners is greatly reduced. The splice plate on one side could be made from the panel itself thus eliminating further rows of fasteners.

As you have it shown, you have rivets in single shear and that is not acceptable for primary structure. There would have to have an internal and/or an external butt strap to add structural integrity. A rivet through three layers of material (double shear) has over twice the shear strength as a rivet through two layers (single shear).


User currently offlineRedFlyer From United States of America, joined Feb 2005, 4335 posts, RR: 28
Reply 12, posted (6 years 8 months 1 week 1 day 7 hours ago) and read 23921 times:



Quoting WingedMigrator (Reply 10):
This link from 2003 has a fascinating, detailed and illustrated description of the manufacturing steps for the A380 bulkhead.

Thanks. That link worked...and it is a terrific article. I loved this part of the article on the use of foam strips...

Quote:
The stiffeners are made with Cytec prepreg (the same material used for the doublers) wrapped around A-profile-shaped Rohacell polymethacrylimide (PMI) foam...

machines the foam stiffeners to the correct dimensions...

The foam acts as a "flying tool," according to Gralfs. "To achieve the desired shape of the stiffeners, we use the foam as a tool or mandrel. It doesn't have a structural function and it stays in the part, since we can't remove it," he explains. The foam's resistance to high temperature and creep allows it to undergo the second autoclave cure cycle, which is necessary to cure the stiffeners, while its light weight doesn't affect part performance.

I thought this was a very profound comment from the 2003 article...

Quote:
"The new SAERTEX plant is no more than 50m behind our fence," says Jens Gralfs, responsible for composite technology development at Stade. "It really simplifies the logistics to have a supplier located so close."

Isn't that what Mike Bair of Boeing is proposing for future manufacturing?



My other home is a Piper Cherokee 180C
User currently offlineWingedMigrator From United States of America, joined Oct 2005, 2214 posts, RR: 56
Reply 13, posted (6 years 8 months 1 week 1 day 7 hours ago) and read 23923 times:



Quoting RedFlyer (Reply 12):
I loved this part of the article on the use of foam strips...

 checkmark  It's precisely what prompted my post on A350 panels back in June. From my layman's perspective, it sounds as if stringers, frames, door and window frames and other localized reinforcements could potentially be manufactured this way.

I wonder what else they've come up with between 2003 and 2008...  scratchchin 


User currently offline474218 From United States of America, joined Oct 2005, 6340 posts, RR: 9
Reply 14, posted (6 years 8 months 1 week 1 day 6 hours ago) and read 23919 times:



Quoting Rheinwaldner (Thread starter):
Forecast 6: The first 787 write off will be after an incident smaller than BA038 because of unrepairable fuselage damage (something like this:

View Large View Medium

View Large View Medium
Click here for bigger photo!

Photo © Jarett Sirko


Photo © Jarett Sirko

, such a hole can not be patched)

Then how do you explain this photo, taken three (3) years later. Almost anything is repairable if you want to spend the time and money.


View Large View Medium
Click here for bigger photo!

Photo © John Padgett



User currently offlineRheinwaldner From Switzerland, joined Jan 2008, 2241 posts, RR: 5
Reply 15, posted (6 years 8 months 6 days 18 hours ago) and read 23817 times:

 Cool Thank you all very much for the contributions to this thread. I appreciate any thought! I knew that some of my thoughts and especially the forecasts were a little provocative.
I do not build aircrafts (though I am an engineer) and many things discussed are admittedly very uncertain. I can not "prove" the points I made and it is easily possible that my conclussions are miles away from how it will turn out "in reality". I also admit that certainly many of you have more expertise than me on the subject. That some of my points receive acceptance pleases me.

Quoting RedFlyer (Reply 7):
in my view there are too many variables and unknowns that prevent the categorical statement that one or the other is better.

Agreed!

But still I am very curious about the discussed technologies and how CFRP fuselages in the future will be build.

I would like to comment some statements:

Quoting Tdscanuck (Reply 4):
There is no particular reason you can't do a barrel with a tooling surface on the outside, although I don't believe the 787 is doing so. I don't think this is an inherent panel vs. barrel issue.

That would mean a mandrel like a tunnel. I can hardly imagine that the fibers would stick at the roof. But if it would work it would allow an interressting process: The mandrel would also serve as the autoclave's case. After moulding a portable heating-unit would be slided into the barrel to bake the form.

Quoting Tdscanuck (Reply 4):
Both panel and barrel construction can support higher integration than either Airbus or Boeing are currently considering. I'm don't see why your point above applies to panels and not barrels.

I think this is true and a mindful observation. But I think too the limit of what can be done is reached earlier with barrels. For barrels any complex structure would require cutouts in the mandrel and the shape would lay in the form. The first steps in the process (layout of the prepreg) become more complicated and also a steady heat distribution is more difficult (if not impossible). The same can be said for the removal of the mandrel.

About manufacturing:

Quoting Tdscanuck (Reply 4):
Absolutely true on a part basis. For an overall airplane cost basis, I doubt it because of the savings in final assembly joints.



Quoting Tdscanuck (Reply 4):
Highly doubtful given that, no matter how you slice it, panels will have more major joints.



Quoting RedFlyer (Reply 5):
Wouldn't the Panel Max approach actually increase manufacturing costs



Quoting RedFlyer (Reply 7):
That was my whole point: "partially curing the parts individually...then do the final co-cure" would seem to be multiple curings. All of which adds time and complexity to the process. It certainly doesn't make it more efficient. The individual parts will still have to be "constructed" prior to the final sub-assembly. Rather than construct the support structures on the back-end of the process as Boeing does on its Barrrel method, Airbus would construct the support structures on the front-end of their Panel method.



Quoting RedFlyer (Reply 9):
No doubt Airbus has experience with this and they could do it very efficiently. What I'm curious to know is how this specifically might be a better (read more efficient) process than Boeing's barrel approach. As I said, Boeing is constructing the stringers, frames, door frames, and window frames on the back-end of the construction process. So Airbus does it on the front end - how does that make it more efficient? Parity, I can understand. I'm not sure I grasp the "more" efficient aspect of it.

Even if only stringers and frames can be co-cured and are completely finished after the autoclave I am quite shure that the manufacturing is leaner. Why:

  • Comparing with the 787 barrels you need also to consider the effort to build the frames. That means for a 10 m long barrel more than 10 parts in addition that need to be manufactured. If they are done with CRFP technology that means time-consuming moulding and baking too.
  • The process to build the co-cured panel with stringers and frames may be time consuming (perhaps requiring severall curings) BUT it is nearly fully automated.
  • In contrast adding the frames manually requires:

    • Huge amount more drilling (possibly automated) but drilling is something you want to avoid as much as possible when dealing with CFRP
    • Manual assembly of the frames (time and labor consuming)
    • Huge amount more fastener

  • An autoclave that takes one barrel probaly can take 6 panels
  • Take a fuselage sector with arbitrary length and count the fasteners either for (a) the frames inside of a barrel or (b) the longitudinal joints for joining four panels. (a) has the higher amount of fasteners.


Thank you again for your interest. I am away from the internet for today.

Regards Martin


User currently offlineTdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 16, posted (6 years 8 months 6 days 12 hours ago) and read 23750 times:



Quoting RedFlyer (Reply 7):
Quoting Tdscanuck (Reply 6):
Quoting RedFlyer (Reply 5):
Wouldn't one also require a unique form or press for each panel?

Unique tooling or reconfigurable tooling. Probably no press involved (that's what the autoclave is for).

Just curious, but how would the parts be held in place during the autoclave process? I was under the impression an autoclave applies pressure and heat to "cure" the compound by removing all air from the material. How would the separate parts be applied (all at the same time, I presume) under pressure while ensuring they remain static?

Removing air is a little different than autoclaving. You can cure a composite part outside an autoclave by pulling a vacuum on it, squashing the part against the tool. Then you just need an oven or heat blanket. This is fairly common way to do repairs. An autoclave squashes the part by pressuring up the autoclave. For the best results, you do both. In all the cases, the "press" is air pressure, not a physical press.

Quoting RedFlyer (Reply 7):
Rather than construct the support structures on the back-end of the process as Boeing does on its Barrrel method, Airbus would construct the support structures on the front-end of their Panel method.



Quoting RedFlyer (Reply 9):
Boeing is constructing the stringers, frames, door frames, and window frames on the back-end of the construction process.

I must be missing something...Boeing does the stringers at the front-end of the process.

Quoting Rheinwaldner (Reply 15):
Quoting Tdscanuck (Reply 4):
There is no particular reason you can't do a barrel with a tooling surface on the outside, although I don't believe the 787 is doing so. I don't think this is an inherent panel vs. barrel issue.

That would mean a mandrel like a tunnel. I can hardly imagine that the fibers would stick at the roof. But if it would work it would allow an interressting process: The mandrel would also serve as the autoclave's case. After moulding a portable heating-unit would be slided into the barrel to bake the form.

The idea of using the mandrel as the autoclave case is slick, I like that. As for having the fibers stick to the roof, I'd just tip the whole thing on end and build it as a vertical cylinder.

Tom.


User currently offlineFlipdewaf From United Kingdom, joined Jul 2006, 1574 posts, RR: 1
Reply 17, posted (6 years 8 months 6 days 10 hours ago) and read 23736 times:
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Quoting Tdscanuck (Reply 16):

The idea of using the mandrel as the autoclave case is slick, I like that. As for having the fibers stick to the roof, I'd just tip the whole thing on end and build it as a vertical cylinder.

In my head I can see that slumping down, what about a spinning autoclave?

Fred


User currently offlineTdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 18, posted (6 years 8 months 6 days 9 hours ago) and read 23730 times:



Quoting Flipdewaf (Reply 17):
Quoting Tdscanuck (Reply 16):

The idea of using the mandrel as the autoclave case is slick, I like that. As for having the fibers stick to the roof, I'd just tip the whole thing on end and build it as a vertical cylinder.

In my head I can see that slumping down, what about a spinning autoclave?

I like it. Although, to be really fancy, we ought to spin it on at least two axes, like rotomolded plastic, to assure even forces on the fibers. Cue the giant spherical autoclave!

Tom.


User currently offlineRheinwaldner From Switzerland, joined Jan 2008, 2241 posts, RR: 5
Reply 19, posted (6 years 8 months 5 days 20 hours ago) and read 23687 times:

I have some more thoughs, remarks and pictures!

About manufacturing

There are several concepts that aim to eliminate the autoclave. One idea is to heat the structures with microwaves ( http://www.dlr.de/fa/Portaldata/17/R...okumente/institut/2005/2005_02.pdf )


About the joints:

Quoting 474218 (Reply 11):
As you have it shown, you have rivets in single shear and that is not acceptable for primary structure. There would have to have an internal and/or an external butt strap to add structural integrity. A rivet through three layers of material (double shear) has over twice the shear strength as a rivet through two layers (single shear).

See this picture:

Do I understand you correctly that c) would be the best? But I think a) is how metal panels are joined today and thus b) would be better than a) because it has only half of the drilling, fasteners and handwork. The area in the green circle does handle forces better in every aspect than the area in the red circle and it still uses less material (-> less weight).

Quoting Tdscanuck (Reply 4):
There's a slight issue here with your lap joint...tremendous off-center loading (and a heck of a stress riser in the sharp corner). Although I certainly agree you could integrate the lap joint into the panel, it strongly doubt the cross section would look like this.

I made two other proposals how efficient joints could look like. Both take advantage of easier shaping of CFRP skins. The solution could be applied to both panels or barrels. Both solutions may eliminate the drilling. This could be achieved by CFRP-ducts that are put in place.

Further reading stuff: http://www.dlr.de/fa/Portaldata/17/R...okumente/institut/2005/2005_03.pdf
and http://www.dlr.de/fa/Portaldata/17/R...okumente/institut/2004/2004_09.pdf


Repairability

Quoting 474218 (Reply 14):
Then how do you explain this photo, taken three (3) years later. Almost anything is repairable if you want to spend the time and money.

This aircraft is THE bechmark for CFRP-repairability!
I am glad that this plane flies again and I did not doubt it. I think repairing this damage is indeed a benchmark for CFRP-repairability. For metal planes it should be easier because every damaged sheet can be reconstructed and be bolted on the frame again. For a CFRP-barrel-fuselage it is IMO out of question that the barrel can economically be replaced as a whole. The procedure to fix the damage had to be something like this:

  • More material had to be cut out (make a clean "rectangular" hole). This chip removing process is not "easy" with CFRP
  • A new sheet would have to be made. This requires first a Mandrel just for this sheet (because Boeing has not the mandrels right now to construct parts of the belly as panels), then moulding and curing. Conclusion: such a piece could probably only be made by Boeing themselves. Also adding the stringers in the same manner as the barrel has them is not easy.
  • Then with joints like a) in the picture above the new panel would be put in place.

I can not imagine that such a procedure will be considered as realistic option in such cases (I know they are seldom anyway). Therefore my forecast: The first 787 write-off will be after an incident like this that could have been repaired with a metal fuselage.
Regarding panels: I addmit the effort to repair such a damage would be tremendous too. It would also need the original panel coming directly from the Airbus factory. Advantages that remain: (a) No different tooling for the "spare-part"-manufacturing required and (b) no chip removing process involved.

About crashworthiness

Quoting Tdscanuck (Reply 4):
Again, not sure why. One of the huge advantages of composites, in both panel and barrel form, is local tailoring of structural properties. You can build a barrel with very different upper and lower crash response just as you can build it with panels.

The core of the difficulty lies in the general structure (semi-monolithic tube) we have today. This conclusion bases on the quote from NASA as cited in the first post:
"The ring-frame configuration as commonly used in fixed-wing aircraft, is a difficult component for crashworthiness. NASA studies have shown, that the "point-load" applied by the ground leads to immediate fracture at this point, followed by severe bending, and further breakages of the frame higher up [7]. This may result in an early disintegration of the structure, and the bypassing of the dedicated energy-absorbing measures. The recommendations followed from these studies were, to separate the livable volume "on top" from the energy absorbing components "below". Hence, a sufficiently strong, closed ring-frame, meant to survive the impact, should be positioned on top of an expandable energy absorbing structure....". (The link to the original source is also in the first post).
The Gondolaconcept as described in this document http://www.dlr.de/wf/en/Portaldata/2...t-fuer-einen-cfk-flugzeugrumpf.pdf (page 11 + 12) and more in detail in this document http://www.dlr.de/fa/PortalData/17/R...blikationen/2004/11_kolesnikov.pdf (from page 8) is a direct one-to-one realization of that NASA thesis. It therefore allows to build in superior crash protection.

The Gondolaconcept
Please read the document from second link (11_kolesnikov.pdf) in the paragraph before. It stresses following points (these are not my opinions I just summarize what is in the document):

  • Page 3: A fuselage is not a tube. Cutouts are exactly there where the bending moments are the largest. Thus the ideal method to make a tube is not nessearily the best method to make an aircraft fuselage. The documents describe this as "Hence the standard fuselage in the lower panel area is not an optimal 'light' construction from the point of view of structural mechanics, and arrangement of pits and cutouts in load carrying structure requires the increased material consumption.". BTW on a sidenote: This tramway was ordered 1996 as 'crude'-winding CFRP monocoque.

    BUT the cutouts of the big windows destroyed the business case. It is now realized with Alu-Hybrid-technique. Nevertheless it carried me well this morning!
  • From Page 8 on: The basis idea is to split the fuselage vertically. If you take away the belly from the monolithic fuselage you get (a) a much more undisturbed monolithic upper part that shall survive crash impacts (b) you can construct a belly has no integral load bearing function but suits much better the relevant requirements.
  • The upper part would IMO be a good candidate to be made with barrels
  • The crash worthiness promises to be much more manageable (according to NASA).
  • The gondola surface allows maintenance reduction by following idea: It is quite weak and if it lacks of any visible buck it is almost guaranteed that no impact occured thus allowing to reduce inspection activity.
  • Increased repairability of the belly is very likely
  • A full scale demonstrator was build as early as 2002. The construction of the demonstrator is described here http://www.dlr.de/fa/Portaldata/17/R...okumente/institut/2003/2003_03.pdf

Further reading stuff: http://www.dlr.de/fa/Portaldata/17/R...okumente/institut/2004/2004_03.pdf


Foam application

Quoting WingedMigrator (Reply 13):
It's precisely what prompted my post on A350 panels back in June. From my layman's perspective, it sounds as if stringers, frames, door and window frames and other localized reinforcements could potentially be manufactured this way.

May I draw your attention to page 9 (upper half) of this document:
http://www.dlr.de/wf/en/Portaldata/2...t-fuer-einen-cfk-flugzeugrumpf.pdf

On that page is a concept shown where foam is used for the very topic of this thread (CFRP fuselage with integrated, intersection-free stringers and frames).

Regards Martin


User currently offline474218 From United States of America, joined Oct 2005, 6340 posts, RR: 9
Reply 20, posted (6 years 8 months 5 days 18 hours ago) and read 23646 times:



Quoting Rheinwaldner (Reply 19):
See this picture:

Do I understand you correctly that c) would be the best? But I think a) is how metal panels are joined today and thus b) would be better than a) because it has only half of the drilling, fasteners and handwork. The area in the green circle does handle forces better in every aspect than the area in the red circle and it still uses less material (-> less weight).

A and B still have rivets in single shear and therefore are structurally unacceptable (for primary structure).

Look at the following site, it describes in detail the effects of having primary structure in single shear. See Figure 5.


http://shippai.jst.go.jp/en/Detail?fn=0&id=CB1071008


By the way, I am not criticizing your work I think what you have done is outstanding.


User currently offlineSoon7x7 From , joined Dec 1969, posts, RR:
Reply 21, posted (6 years 8 months 5 days 16 hours ago) and read 23637 times:



Quoting 474218 (Reply 14):

Composite aircraft are throw away birds, metal fabricated aircraft are easily repairable, are predictable in terms of corrosion life, frame failures, (gives clues to impending failures) composites typically just fail catastrophically, I feel with this new technology, a whole new learning curve will be realized with aircraft already in revenue service at a cost...Torque boxes and major airframe structures fabricated from plastics makes no sense to me. Wing tips, fairings, and other associated streamlining panels make complete sense. Composites despite how the industry boasts about it benefits, are also subject to the elements if not protected and constantly maintained. Same can be said about metal fabrication but alloy structures have proven over time to be unequalled in durability, just take a walk through the boneyards out west, Fourty, sixty year old metal birds still sitting out there in potentially flying condition, yet....scout around the yards for composite parts and they exhibit great deterioration from the natural elements.I'm speaking in general terms as arguments can be made both pros and cons regarding both methods but I feel the new technology with ultimetelly prove to be a white elephant...


User currently offlineRedFlyer From United States of America, joined Feb 2005, 4335 posts, RR: 28
Reply 22, posted (6 years 8 months 4 days 6 hours ago) and read 23546 times:



Quoting Tdscanuck (Reply 16):
Removing air is a little different than autoclaving.

I meant removing air from the material being cured. Isn't that what the big concern was at Boeing with making the first barrels? They had to ensure there were no air pockets/bubbles in the CFRP. Or did I misunderstand what you were saying?

Quoting Tdscanuck (Reply 16):
Boeing does the stringers at the front-end of the process.

Yes, I know. I was trying to keep the analysis simple by showing that both methods use sub-structures, which require manufacturing prior to mating to the sub-assembly structures. I was just trying to keep the analogies simple so as not to confuse.  Smile

Quoting Soon7x7 (Reply 21):
Composite aircraft are throw away birds, metal fabricated aircraft are easily repairable, are predictable in terms of corrosion life, frame failures, (gives clues to impending failures) composites typically just fail catastrophically, I feel with this new technology, a whole new learning curve will be realized with aircraft already in revenue service at a cost...Torque boxes and major airframe structures fabricated from plastics makes no sense to me.

My first reaction is to say "get used to it because composites, be they the Airbus way or the Boeing way, are coming to an airport near you soon!"  Wink But I think the fact is that there is a lot of confidence in metal fabricated aircraft because we've had about 80 years of experience with them, 50 in the jet age alone. That's no excuse to stop progress. I'm sure the same arguments could have been made when manufacturers gave up wood and fabric for the first aluminum birds.



My other home is a Piper Cherokee 180C
User currently offlineTdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 23, posted (6 years 8 months 4 days 4 hours ago) and read 23535 times:



Quoting RedFlyer (Reply 22):
Quoting Tdscanuck (Reply 16):
Removing air is a little different than autoclaving.

I meant removing air from the material being cured. Isn't that what the big concern was at Boeing with making the first barrels? They had to ensure there were no air pockets/bubbles in the CFRP. Or did I misunderstand what you were saying?

Ah, gotcha. Autoclaving alone will reduce the size of voids but not eliminate them. For that you need to vacuum somehow. I thought you just meant removing air around the part so that the bag would press down on it.

Quoting Soon7x7 (Reply 21):
Composite aircraft are throw away birds, metal fabricated aircraft are easily repairable, are predictable in terms of corrosion life, frame failures, (gives clues to impending failures) composites typically just fail catastrophically, I feel with this new technology, a whole new learning curve will be realized with aircraft already in revenue service at a cost...

"New technology"? Airbus has had composite primary structure since the 80's. It's been tested on aircraft since at least the 70's. Coupon-level testing has studied the corrosion and fatigue life out for many decades.

Quoting Soon7x7 (Reply 21):
Torque boxes and major airframe structures fabricated from plastics makes no sense to me.

Except it's been going on for ages. Vertical stabilizers, thrust reversers, wing panels, floor beams, fan blades, wings...the list goes on.

Quoting Soon7x7 (Reply 21):
Same can be said about metal fabrication but alloy structures have proven over time to be unequalled in durability

Not really...the durability record on major CFRP parts like A300 vertical stabilizers, 777 floor beams, or CFRP fan blades is *far* better than their alloy counterparts.

Tom.


User currently offlineRheinwaldner From Switzerland, joined Jan 2008, 2241 posts, RR: 5
Reply 24, posted (6 years 8 months 3 days 18 hours ago) and read 23493 times:



Quoting Soon7x7 (Reply 21):
are predictable in terms of corrosion life, frame failures, (gives clues to impending failures) composites typically just fail catastrophically, I feel with this new technology, a whole new learning curve will be realized with aircraft already in revenue service at a cost

I tend to disagree. Currently there is a new generation of aluminium-alloys available too that promises to equal the weight and durability of CFRP BUT they are not one bit better regarding the points you raised. In other words: perdictability about corrosion is not better and the learning curve exists too.
Further improved CFRP materials, that will appear in the future, will offer strengths that can no longer be achieved by aluminium.


25 RedFlyer : I thought CFRP materials already do that - offer a strength to weight ratio that betters that found with metal alloys?
26 WingedMigrator : I'll let others more knowledgeable comment in detail, but research on better alloys is not standing still. The search continues for lighter, stronger
27 Post contains links Rheinwaldner : This document on page 5 below indicates what I have written: http://www.dlr.de/wf/en/Portaldata/2...t-fuer-einen-cfk-flugzeugrumpf.pdf The horizontal
28 RedFlyer : I thought I'd re-light this very interesting thread as I don't believe there has been any new discussions on Airbus' panel approach to construction of
29 Post contains links Rheinwaldner : Hi I had some time to find these links again but here they are (it was discussed in this thread http://www.airliners.net/aviation-fo...general_aviati
30 Soon7x7 : Another aspect of carbon fibre technology is crash survivability...not in the human sense but technically speaking, airframe survival...Roselawn, Indi
31 Parapente : What afantastic post -and replies! with such a high quality of discussion I will not attempt to comment meaningfully on that which I am not qualified
32 Post contains links Keesje : I think the loads on e.g. the upperside of the fuselage are not the same as in the side and lower parts of the fuselage. Panels can be optimized for
33 Tdscanuck : It's absolutely true that the loads are different around the circumference, but there's nothing that prevents you from doing the same thing with the
34 Post contains images Keesje : I think discontimued layers create problems of their own. The tape laying machines cuts off the prepreg tape 90 degrees angled of the tape laying dir
35 Tdscanuck : It's done all the time...it's really not a big deal. There are discontinuous edges in all laminates of anything more than trivial layup. If you requi
36 RedFlyer : Flag waving, regardless of the flag, is good. Wave yours proudly. Keesje, isn't what you describe here also a problem for panels where, unlike on a c
37 Post contains links RedFlyer : I found this interesting from the referenced article: [emphasis added] Also... [emphasis added] http://www.marketwatch.com/news/stor...-4E9D-AF79-8F60
38 WingedMigrator : Except Boeing uses an inner (convex) mandrel, and Airbus will use an outer (concave) mandrel if I'm not mistaken. That's interesting. I thought the 7
39 Tdscanuck : I believe it's a mix...I think the doors are laid right in but the windows are cut out afterwards. Tom.
40 Rheinwaldner : That is a very nice source. This paragraph reveals even more: "The multi-strand control allows wrinkle-free, near-net shape lay-up of enclosed and de
41 Post contains images Keesje : I don't think so. Is it correct that the fuselage has a constant skin thickness / fiber directions are the same everywhere?[Edited 2008-09-22 08:06:2
42 Tdscanuck : I think the outside mandrel gives you more flexibility (you can change the layup without changing the tool) but I'm not sure it limits you in any par
43 Post contains links and images Rheinwaldner : The first point IMO you mention is already significant. For the 787 does that mean the 783, 788, 789, 781 get all different mandrels? Otherwise how t
44 Tdscanuck : I'm honestly not sure. Given expected production rates, they might. There's a couple of things they can do...the easiest (probably) is to tweak the s
45 Rheinwaldner : I know that Boeing has accomplished to integrate the stringers in the panels before curing! This is IMO quite an achievement and I doubt that much mo
46 Pianos101 : Yeah it's definitely not constant. You don't even need to see the ply map, you can just look at the solid model in catia... I'm at Vought now and you
47 Tdscanuck : All of the illustrations you showed could be done with an inside mandrel (don't forget that, in the second photo, the engine mounts are put on later)
48 Astuteman : Great thread this guys. Keep up the good work. Mind you, it's very entertaining reading the old threads linked in the OP. My my. How things can change
49 Rheinwaldner : That happens to be true until now, but with outside mandrels such changes may become cheap. Local variations of the fuselage shell strength is not mo
50 Pianos101 : At least you understand the last point. This is not even remotely possible. On the software side, yeah it would probably be easy, but where does the
51 Tdscanuck : I wasn't very clear...the big payment wouldn't be because of the tooling and programming...as you correctly note, that would be pretty minor. With a
52 Pianos101 : It is a uniform gauge, but it's still theoretical. You can't just cut holes willy-nilly through the skin (even with avoiding the frames) without unde
53 Areopagus : The photo doesn't give details on the collision, but Boeing has said in the past that CFRP is more resilient, and not as likely to be damaged by ramp
54 Post contains links RedFlyer : http://www.reuters.com/article/rbssI...sUtilitiesNews/idUSL84625120081008 While a relatively small sum at the moment, I presume this is for R&D of th
55 Post contains links and images Rheinwaldner : Informations about the next CFRP fuselage project are in the news: http://www.aviationweek.com/aw/blogs...918845-aaa6-42f3-adb8-895b2b26d513 Interesti
56 RedFlyer : The traditional large autoclave is dismissed, but this comment... ...seems to imply that individual autoclaves are built around each component. Or am
57 SEPilot : From what I have read about the panel vs. barrel argument, plus my own experience as a design engineer, I am convinced that the barrels are superior a
58 Tdscanuck : That shouldn't be surprising to anyone...the barrel approach has lots of advantages, but cheapness isn't one of them (especially capital outlay). As
59 Astuteman : That's probably one of the most insightful comments I've ever read on here, Tom, in every phrase Everything in our world is a trade off of several fa
60 RedFlyer : That brings up an interesting point: if this next advance in CFRP construction has resulted in elimination of "pressure" from the cure process, will
61 Rheinwaldner : Thanks for engineering input! Is always welcome! I think your thoughts are valid. But on the other hand I am not sure whether the focus is not too mu
62 SEPilot : I was not aware of this one; I will have to look it up. Of course a fuselage is more than just a pressurized tube, but that is where by far the great
63 Post contains links Rheinwaldner : I agree fully about the key focus on performance (=payload/range figures + efficiency). The design of airliners does stress the performance objective
64 SEPilot : Parts count is not necessarily indicative of performance (although reducing it is almost always desireable.) In the case of barrels vs. panels the jo
65 Post contains links Tdscanuck : I'm sure there are lots of resin engineers thinking hard about that question. It's actually doable today, but not without so severe a performance hit
66 JoeCanuck : One of the problems with chemically cured epoxies is handling time. You can do very large pieces and have them cure uniformly using heat/pressure as
67 SEPilot : " target=_blank>http://faalessons2006.workforceconne..._Air/ I had already found this; I read the whole site. Fascinating material, and it goes to sh
68 474218 : Not all: The Lockheed L-1011 has no 'hydraulic lines' in the horizontal stabilizer.
69 Tdscanuck : How were they moving the elevators? Tom.
70 474218 : The elevators are slaved off of the flying stabilizer. When the stabilizer nose goes down the elevators are moved up by 3/8" diameter cables. When th
71 RedFlyer : But in a production environment where every process is meticulously choreographed, wouldn't/shouldn't delays in lay-up be a minor risk? As far as I c
72 JoeCanuck : The pieces both Airbus and Boeing are dealing with are huge. They require a great deal of time to assemble so it would be almost impossible to time t
73 Rheinwaldner : Glad you noticed the discrepancy. The problem is only that IMO the A350 is "in line" with what can be expected from CFRP and the 787 is not. There a
74 SEPilot : If that is indeed the case, I cannot imagine that Boeing will be doing a major weight reduction at some point on the 787. I have not looked at the nu
75 JoeCanuck : I remain unconvinced, I'm afraid. There is no guarantee that installing windows and other structure in the manner you visualize will net any signific
76 RedFlyer : R, Airbus is at the development point of their project that Boeing was when they were claiming weight efficiencies. Will this paradox still hold true
77 Tdscanuck : Absolutely. I hope Airbus pulls it off (it will be a marvelous piece of structural engineering if they do). I'm just confused how they're going to do
78 Astuteman : Although they've been building large CFRP panels for a long time now... Curiously, Airbus did just add in 2.2 tonnes to cover the same lightning prot
79 Brendows : Just the skin, that is excluding stringers, frames, etc? Then I'd guess that the figure is below 15%. A couple of years ago I saw some weight figures
80 SEPilot : Excellent observation; I have come to the conclusion that fuselage diameter is a hugely important factor in efficiency. I think this is the biggest r
81 Nomadd22 : The two factors that I can think of to account for that would be Airbus having more confidence in their projections and/or Airbus being willing to ta
82 RedFlyer : Good observation, Astute. In the final analysis, we just never have any idea if when we're kicking around these numbers we're really making an apples
83 Astuteman : I can buy that. I posed the question to contextualise the impact of migrating to CFRP on the overall aircraft weight, and perhaps examine why some of
84 SEPilot : Totally agree. But there aren't enough of us to say that we'll pay extra to go on a 767 over an A330. That is why the airliner designers get the big
85 Nomadd22 : Last night I had a dream in which God revealed to me that henceforth, all seating comparisons will be made by however many 17.2" seats will fit in a
86 RedFlyer : Okay, Av Fans, time to resurrect this splendid little thread. In light of recent developments at Boeing with regards to their stress-test failure poin
87 Astuteman : Agree Although they already have the worlds largest CFRP wingbox in airline service (probably the worlds largest wingbox full-stop), not to mention t
88 WingedMigrator : It may also have to do with interface management. The location of the problem is where four sections made by four different companies, each with thei
89 Rheinwaldner : I agree with that. In fact you could even argue that the 787 center fuselage by its nature is more panel technology than barrel technology. At least
90 Keesje : I can imagine Airbus opting for less controversial solutions if they have a choice. Conservatives gaining power. Maybe the amount of composites will
91 RedFlyer : They are as I think the stringers are CF in nature for both models with the frames being metal. But using Rheinwaldner's assumptions in his original
92 WingedMigrator : Not sure what you mean here. Regardless of manufacturing technique, the end result is the same: a semi-monocoque structure with frames, stringers and
93 NicoEDDF : Uhh...ugly prediction from March this year, Nomadd!
94 RedFlyer : What you wrote is exactly what I was saying: The end result is the same; just the method of getting there is different.
95 Tdscanuck : That's not how this reads: Since the end result is the same, there's no theoretical advantage for Airbus in load transfer. Tom.
96 RedFlyer : I believe that my comment that you quoted was regarding obtaining the outcome only, that being a semi-monocoque structure. And even at that, there co
97 Castillo : In general, whenever you introduce a bolted load-transfer detail into a structure you reduce the stress that you can work the structure to, which tra
98 Tdscanuck : It's not the same semi-monocoque structure. They're both semi-monocoque structures, but that's the principal, not the realization. A Douglas fuselage
99 Rheinwaldner : Does that mean that if frames and stringers could be co-cured in one step (as per my initial idea) there would be some savings? And would barrels wit
100 Tdscanuck : Yes. Definite reduction in fasteners, if nothing else. Yes. In theory, as far as panels. It's a tooling problem, not a part problem. It's all about t
101 Post contains links Rheinwaldner : Concerning the recent 787 delay and the barrel approach this little information emerged: "That may involve reworking composite molds" -> from here htt
102 RedFlyer : We can also say that Boeing's methodology has caused them considerable problems, as evidenced from the fact that their airplane is heading into a 3 y
103 WingedMigrator : To be fair, the "reworking" of wing / wing box components described in this article does not involve fuselage barrel structures... therefore, the who
104 Tdscanuck : What, of the current or past issues, had anything to do with barrels vs. panels? I'm certainly not claiming the 787 doesn't have problems, just that
105 Rheinwaldner : I agree we can't say for sure. But it could be that because of the body-wing-join problem the need could arise to enforce the fuselage skin at those
106 Tdscanuck : I agree that part change is easier on an inside mandrel. However, I'm not sure how much of an obstacle tool change is. To recontour a mandrel is anno
107 Post contains links Rheinwaldner : I borrowed this research report from the ETH (Swiss technical university) library: Link to library The following highlights summarize this research re
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