rheinwaldner
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Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 10:17 am

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).
 
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autothrust
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RE: Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 1:47 pm

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
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hawkercamm
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RE: Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 2:51 pm

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
 
Arniepie
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RE: Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 3:26 pm



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]
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 4:54 pm



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.
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 7:22 pm

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.
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tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Fri Jan 25, 2008 8:31 pm



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.
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Sat Jan 26, 2008 3:55 am



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.
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WingedMigrator
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RE: Cfrp Panel Fuselages Superior (?)

Sat Jan 26, 2008 5:04 am

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]
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Sat Jan 26, 2008 7:57 am

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]
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WingedMigrator
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RE: Cfrp Panel Fuselages Superior (?)

Sat Jan 26, 2008 6:18 pm



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.
 
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RE: Cfrp Panel Fuselages Superior (?)

Sat Jan 26, 2008 7:44 pm



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).
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Sun Jan 27, 2008 3:20 am



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 in the sky inside my Piper Cherokee 180.
 
WingedMigrator
Posts: 1769
Joined: Wed Oct 26, 2005 9:45 am

RE: Cfrp Panel Fuselages Superior (?)

Sun Jan 27, 2008 3:29 am



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 
 
474218
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RE: Cfrp Panel Fuselages Superior (?)

Sun Jan 27, 2008 3:47 am



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

 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Mon Jan 28, 2008 3:59 pm

 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
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Mon Jan 28, 2008 10:29 pm



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.
 
flipdewaf
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RE: Cfrp Panel Fuselages Superior (?)

Tue Jan 29, 2008 12:08 am



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
Image
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Tue Jan 29, 2008 1:25 am



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.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Tue Jan 29, 2008 1:42 pm

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
 
474218
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RE: Cfrp Panel Fuselages Superior (?)

Tue Jan 29, 2008 4:04 pm



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.
 
soon7x7
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RE: Cfrp Panel Fuselages Superior (?)

Tue Jan 29, 2008 5:51 pm



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...
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Thu Jan 31, 2008 4:19 am



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 in the sky inside my Piper Cherokee 180.
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Thu Jan 31, 2008 5:40 am



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.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Thu Jan 31, 2008 3:57 pm



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.
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Fri Feb 01, 2008 1:42 am



Quoting Rheinwaldner (Reply 24):
Further improved CFRP materials, that will appear in the future, will offer strengths that can no longer be achieved by aluminium.

I thought CFRP materials already do that - offer a strength to weight ratio that betters that found with metal alloys?
My other home is in the sky inside my Piper Cherokee 180.
 
WingedMigrator
Posts: 1769
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RE: Cfrp Panel Fuselages Superior (?)

Fri Feb 01, 2008 3:51 am



Quoting RedFlyer (Reply 25):
I thought CFRP materials already do that - offer a strength to weight ratio that betters that found with metal alloys?

I'll let others more knowledgeable comment in detail, but research on better alloys is not standing still. The search continues for lighter, stronger, more fatigue resistant, and more corrosion resistant alloys.

Likewise, there is research beyond just CFRP and alloys... such as fiber / alloy laminates. GLARE is an early example thereof, which may yet develop into something more capable.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Fri Feb 01, 2008 1:12 pm



Quoting RedFlyer (Reply 25):
I thought CFRP materials already do that - offer a strength to weight ratio that betters that found with metal alloys?

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 axis is cost (less to the left) and the vertical axis is weight (less at the bottom).
The red area is for new metal technologies. It matches the area of present CFRP technologies (light blue area). Future CFRP technologies will supersede both in cost and weight (dark blue area).

Regards Martin
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Tue Sep 16, 2008 6:47 pm

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 the A350. Has there been any new developments or news regarding Airbus' panel construction approach? Are they still perfecting their methodology or have they settled on the final specifications? I believe design freeze for the A350 is looming so that would mean the construction specs would also be nearing final stages.
My other home is in the sky inside my Piper Cherokee 180.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Fri Sep 19, 2008 3:20 pm



Quoting RedFlyer (Reply 28):
Has there been any new developments or news regarding Airbus' panel construction approach?

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_aviation/read.main/3995067 ) :
http://www.aero.de/news.php?varnewsid=6502
http://www.flightglobal.com/articles...re-a350-fuselage-demonstrator.html

These are demonstrator parts and they do not show anything new IMO regarding the aspects discussed in this thread. Is there some thing more under the hood? I don't know. Will it turn out that no idea presented in this thread will be implemented? Maybe, still a little early to draw conclusions. I would have expected at least stringers AND frames integrated on the panels.
 
soon7x7
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RE: Cfrp Panel Fuselages Superior (?)

Fri Sep 19, 2008 7:54 pm

Another aspect of carbon fibre technology is crash survivability...not in the human sense but technically speaking, airframe survival...Roselawn, Indiana...ATR crash...investigators were confused when discovering they DISCOVERED no wings...turned out, the CARBON FIBRE wings desintegrated on impact...creating an extremely hazardous environment for investigators and reducing the amount of structural evidence available for investigators to sift through. As far as passenger survival in a crash, now in addition to poisonous gases emited during post crash fires, airborne carbon fibre dust/ boron, etc are introduced into the noxious nexes which, much like the 9/11 building dust has been responsible for dicease and ultimetly death for many exposed. When composites burn out in fire, all that remains is a pile of loose untwined fabric showing no definate shape, no part numbers..nothing for investigators to go on...At least in panel contruction, the general framework may still yield clues to incident/accident event causes.
 
parapente
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RE: Cfrp Panel Fuselages Superior (?)

Fri Sep 19, 2008 8:45 pm

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 to do so.

Perhaps save to say that Boeing now openly admit that early aircraft will be heavier than forecast -indeed with this weeks release of the HGW varient of the 330 Airbus hopes to capitalise on this fact in the short term. What is far more worrying is the complete absence of the aircraft for more than a year. But unlike the 350 it's not paper.

What I will say is this.

How nice for us in the west to be able to have this discussion at all. Surrounded as we are by the collapse of our derivative based finance markets is it not a fantastic thing that there is at least one manufacturing area where such a discussion is posssible. I have an IT company and am proud that this industry too is lead by the same countries.

Yes its flag waving - but we have so little to wave these days!
 
keesje
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RE: Cfrp Panel Fuselages Superior (?)

Fri Sep 19, 2008 9:49 pm



Quoting Tdscanuck (Reply 4):
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.

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 the loads at specific locations, saving weight compared to a more homogeneous barrell.

Quoting WingedMigrator (Reply 8):
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.

I know Thermoplastic materials like used on the A380 J noses and tail are considered for more main structures. They can be welded & produced faster / cheaper without autoclaves.

http://www.compositesworld.com/artic...gain-leading-edge-on-the-a380.aspx
"Never mistake motion for action." Ernest Hemingway
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Fri Sep 19, 2008 10:41 pm



Quoting Keesje (Reply 32):
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 the loads at specific locations, saving weight compared to a more homogeneous barrell.

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 barrel as you do with the panel. Just like the panel, the barrel is laid up in multiple layers, which can be customized around the circumference. There's no layup that you can do in panels that you can't do lighter in barrels (due to the absence of the joint reinforcement and splice plates).

Tom.
 
keesje
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RE: Cfrp Panel Fuselages Superior (?)

Fri Sep 19, 2008 11:15 pm



Quoting Tdscanuck (Reply 33):
Just like the panel, the barrel is laid up in multiple layers, which can be customized around the circumference. There's no layup that you can do in panels that you can't do lighter in barrels (due to the absence of the joint reinforcement and splice plates).

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 direction. Different layers have different directions, imagine the discontious edges and danger of delamination, different cut off layers over each other, air pockets etc. I think it's not done on the 787.

"Never mistake motion for action." Ernest Hemingway
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Sat Sep 20, 2008 12:25 am



Quoting Keesje (Reply 34):
I think discontimued layers create problems of their own.

It's done all the time...it's really not a big deal.

Quoting Keesje (Reply 34):
The tape laying machines cuts off the prepreg tape 90 degrees angled of the tape laying direction. Different layers have different directions, imagine the discontious edges and danger of delamination, different cut off layers over each other, air pockets etc.

There are discontinuous edges in all laminates of anything more than trivial layup. If you require continuous fibers from edge to edge at all layers, the only thing you have control of is orientation, which would significantly detract from the flexibility you get from composites. Delamination at the discontinuous edge isn't a big deal because the stress has to drop to zero at the edge anyway. Air pockets are why you cure in an autoclave in the first place...done right, they're also not a big deal.

Quoting Keesje (Reply 34):
I think it's not done on the 787.

It's done on the 737, 747, 757, 767, and 777...I'd be absolutely stunned if it wasn't done on the 787.

Tom.
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Sat Sep 20, 2008 5:25 pm



Quoting Parapente (Reply 31):
Yes its flag waving - but we have so little to wave these days!

Flag waving, regardless of the flag, is good. Wave yours proudly.  Wink

Quoting Keesje (Reply 34):
I think discontimued layers create problems of their own.



Quoting Keesje (Reply 34):
imagine the discontious edges and danger of delamination, different cut off layers over each other,

Keesje, isn't what you describe here also a problem for panels where, unlike on a continuous surface, you would have an abrupt cut off of all layers at the edges of each panel? Of all the benefits and weaknesses of both methods, I don't think this argument is particularly valid as a benefit to the panel approach. So far, from what I've seen discussed here, I don't think from a quality standpoint either approach is better than the other. However, I'm beginning to believe the Airbus panel approach might be less complicated from a construction perspective, which is to say it may be more cost-effective. Time will tell.
My other home is in the sky inside my Piper Cherokee 180.
 
redflyer
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RE: Cfrp Panel Fuselages Superior (?)

Sun Sep 21, 2008 5:24 pm

I found this interesting from the referenced article:

Quote:
Airbus has placed orders with MAG Cincinnati for six VIPER(R) Fiber Placement Systems (FPS). Announced at the International Manufacturing Technology Show in Chicago, the Vipers will perform automated fabrication of composite fuselage panels for the up-coming composite airframe A350 XWB. The six VIPER 6000 series FPS, MAG Cincinnati's largest, can produce fuselage panels up to 6.3 m / 20.7 ft in diameter.

[emphasis added]

Also...

Quote:
VIPERs enable independent control over feed, clamp, cut and start for up to 32 individual tows or slit tape. This allows automated "on-the-fly" adjustment of fiber band width, controlled bending of fiber layout around changing contours, and precise configuration of openings (doors, hatches, etc.).

[emphasis added]

http://www.marketwatch.com/news/stor...-4E9D-AF79-8F603EB322CC}&dist=hppr

It appears Airbus will be using the same techniques for lay-up of the CFRP material as Boeing will, only on a smaller scale via their panel approach.
My other home is in the sky inside my Piper Cherokee 180.
 
WingedMigrator
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RE: Cfrp Panel Fuselages Superior (?)

Mon Sep 22, 2008 1:58 am



Quoting RedFlyer (Reply 37):
It appears Airbus will be using the same techniques for lay-up of the CFRP material as Boeing

Except Boeing uses an inner (convex) mandrel, and Airbus will use an outer (concave) mandrel if I'm not mistaken.

Quoting RedFlyer (Reply 37):
and precise configuration of openings (doors, hatches, etc.)

That's interesting. I thought the 787 doors, windows and other openings were cut out after the barrel is cured, with frames / reinforcement fastened afterward, from the inside? This sounds different, potentially for something like Rheinwaldner's "panel max" approach.
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Mon Sep 22, 2008 6:12 am



Quoting WingedMigrator (Reply 38):
That's interesting. I thought the 787 doors, windows and other openings were cut out after the barrel is cured, with frames / reinforcement fastened afterward, from the inside?

I believe it's a mix...I think the doors are laid right in but the windows are cut out afterwards.

Tom.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Mon Sep 22, 2008 11:55 am



Quoting RedFlyer (Reply 37):
http://www.marketwatch.com/news/stor...-4E9D-AF79-8F603EB322CC}&dist=hppr

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 deeply contoured structures and concave/convex surfaces, enabling precision production of fuselage sections, panels, cowls, ducts, and nozzle cones for commercial, military and space vehicles."

IMO these are hints that Airbus goes a step further. The core of my initial thought was that a for a fuselage wall you want to have a complete flat outer surface and as much as possible of 3D structure on the inside. Thus if the inside lays in a mandrel (barrel approach) you probably are far more limited in building various structures than with the panel approach. Additionally the outer surface that should be as even as possible surely can be controlled better with an outside mandrel. Two-to-zero for the outside mandrel. Thus I am looking eagerly for such hints that these A350 panels will incorporate a much higher degree of integration.

The goal should be to produce panels fully automated that contain all 3D Structure (frames, stringers, door frames, window frames, crash absorbing elements, ...) from the beginning.
 
keesje
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RE: Cfrp Panel Fuselages Superior (?)

Mon Sep 22, 2008 3:03 pm

Quoting Tdscanuck (Reply 39):
I believe it's a mix...I think the doors are laid right in but the windows are cut out afterwards.

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 08:06:20]
"Never mistake motion for action." Ernest Hemingway
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Tue Sep 23, 2008 12:59 am



Quoting Rheinwaldner (Reply 40):
IMO these are hints that Airbus goes a step further. The core of my initial thought was that a for a fuselage wall you want to have a complete flat outer surface and as much as possible of 3D structure on the inside. Thus if the inside lays in a mandrel (barrel approach) you probably are far more limited in building various structures than with the panel approach.

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 particular way...if you shape the inside mandrel appropriately you can lay it up most any way you want.

Quoting Keesje (Reply 41):
Quoting Tdscanuck (Reply 39):
I believe it's a mix...I think the doors are laid right in but the windows are cut out afterwards.

I don't think so.

You're right...I thought Vought did it differently but, apparently, they do the trim prior to removal of the mandrel. I'd seen pictures of the section still on the mandrel with openings, which is what I was thinking of, but this was post-trim.

Quoting Keesje (Reply 41):
Is it correct that the fuselage has a constant skin thickness / fiber directions are the same everywhere?

I haven't actually seen a ply map, but I'd be stunned if this were the case. Not only would it be a horrendous waste of the potential of CFRP, it would make the use of automated tape-layers very odd...if it's all one thickness/direction the layup wouldn't be nearly that complicated.

Tom.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Tue Sep 23, 2008 8:05 am



Quoting Tdscanuck (Reply 42):
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 particular way...if you shape the inside mandrel appropriately you can lay it up most any way you want.

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 to tailor the wall thickness for the different strengths that are required? Either they have different tools for each subtype (very costly) or they don't care and build just one strength (and loose the very promising possibility to tailor the structure strengths exactly against the requirements). Could explain IMO why the 783 is such a lemon. It flies the same structure that supports a much higher MTOW. The A350 fuselage panel allow to adjust the local strength for each subtype much easier. This is a big advantage for production and production cost. If barrels must be optimized for varying needs the freedom to do so is either much more limited or on the other hand very costly. Each tiny change means new mandrels to be made.

I have explored the internet a little more and have found some interesting pictures. They show what I meant when I said that the outside mandrel allows much more flexibility. These pictures show some examples of approaches that are completely out of scope for an inside mandrel (barrel approach), ever:
...
These pictures come from a website for CFRP fishing boat construction (1991): http://www.fao.org/docrep/003/t0530e/T0530E00.htm#TOC

I think something like this is the limit for an inside mandrel:

...
But look at this. None of these constructing elements will ever happen on an inside mandrel:
BTW I can not see any difference between the solution on this picture and the foamed rips for the A380 bulkhead as discussed in this thread.

...
or this:

...
or this:


These pictures show what I had in mind when I said the outside mandrel allows 3D-constructing elements on the inside. Frames, stringers, hinges, mounting parts all these things could be made one part with the hull itself. The reduction of fastened connections would be paramount.

Amazing too that this UN documents (thought to share know how about fishing boats to the developing world) already shows these ideas. Notice they promote the outside mandrel for ships too.
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Wed Sep 24, 2008 12:29 am



Quoting Rheinwaldner (Reply 43):
For the 787 does that mean the 783, 788, 789, 781 get all different mandrels?

I'm honestly not sure. Given expected production rates, they might.

Quoting Rheinwaldner (Reply 43):
Otherwise how to tailor the wall thickness for the different strengths that are required?

There's a couple of things they can do...the easiest (probably) is to tweak the stringer layup (since that's done off-mandrel). Re-orienting fibers doesn't change thickness, so that may also be an option. Removing a whole ply is also an option.

Quoting Rheinwaldner (Reply 43):
Either they have different tools for each subtype (very costly)

Most of the cost is the tool itself...the delta between types is relatively small. If the production rate is high enough to support multiple mandrel (which I believe it already is, and it's supposed to go higher) there may not be any particular cost penalty.

Quoting Rheinwaldner (Reply 43):
Each tiny change means new mandrels to be made.

Definitely not. You'd certainly go ahead and just change the mandrel (build it up or mill it down) before you went to a whole new tool.

Quoting Rheinwaldner (Reply 43):

But look at this. None of these constructing elements will ever happen on an inside mandrel:

Err...yes, they do. Your second picture, section b), it almost exactly how the 787 fuselage is done (minus the hollow former).

Quoting Rheinwaldner (Reply 43):
Frames, stringers, hinges, mounting parts all these things could be made one part with the hull itself.

787 stringers are one part with the skin. They're co-cured.

Tom.
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Wed Sep 24, 2008 6:48 am



Quoting Tdscanuck (Reply 44):
787 stringers are one part with the skin. They're co-cured.

 checkmark 
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 more is possible in this regard with inside mandrel. Literal translation of two German phrases would say: "end of flagstaff" or "the highest of all feelings".

Quoting Tdscanuck (Reply 44):
Quoting Rheinwaldner (Reply 43):

But look at this. None of these constructing elements will ever happen on an inside mandrel:

Err...yes, they do. Your second picture, section b), it almost exactly how the 787 fuselage is done (minus the hollow former).

Ok, my sentence was overstated. The word "None" should have been "Most ... will never". How about the rest? Could there be some benefit in applying such construction elements for aircraft fuselages? I assume there exist some nice opportunities.

Quoting Tdscanuck (Reply 44):
Most of the cost is the tool itself...the delta between types is relatively small. If the production rate is high enough to support multiple mandrel (which I believe it already is, and it's supposed to go higher) there may not be any particular cost penalty.

I agree that what you mentioned makes it easier to handle inside mandrel vs. changing requirements.
I see another problem: The inside mandrel blocks also easy configuration changes that could arise from customer wishes (and in small numbers). Things like changing the number of windows (have blind windows) or the number of emergency exits and such things can not be delivered by the basic frame. The freighter version would use new mandrels with no windows.
For narrowbody IIRC there are options to choose the number of overwing exits. With an inside mandrel the local reinforcement either will appear on each copy (used or not) ... or make another new mandrel! Such changes sometimes are demanded only for a small number of frames!
 
pianos101
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RE: Cfrp Panel Fuselages Superior (?)

Wed Sep 24, 2008 2:56 pm



Quoting Tdscanuck (Reply 42):
I haven't actually seen a ply map, but I'd be stunned if this were the case. Not only would it be a horrendous waste of the potential of CFRP, it would make the use of automated tape-layers very odd...if it's all one thickness/direction the layup wouldn't be nearly that complicated.

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 can easily see the ply drops on the inside of the skin.

Quoting Rheinwaldner (Reply 43):
For the 787 does that mean the 783, 788, 789, 781 get all different mandrels?

I don't think there's any reason for them to use the same tooling. That said, many sections of the fuselage might not need to be different from the -8. Just adding a plug in the skin might work; i think that's how some of the (metal) derivatives are done. On top of that, as of right now the -9 is a while away, and the -3 and -10 pretty much don't exist at all...
 
tdscanuck
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RE: Cfrp Panel Fuselages Superior (?)

Thu Sep 25, 2008 4:24 am



Quoting Rheinwaldner (Reply 45):
Quoting Tdscanuck (Reply 44):
Your second picture, section b), it almost exactly how the 787 fuselage is done (minus the hollow former).

Ok, my sentence was overstated. The word "None" should have been "Most ... will never". How about the rest? Could there be some benefit in applying such construction elements for aircraft fuselages? I assume there exist some nice opportunities.

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). The only one that would be sort of tricky would be d) in the second photo, but you'd just need a temporary spacer that you removed after curing.

There certainly is benefit to applying these types of elements to aircraft fuselages, and I'm pretty sure you could find examples of all of these in aviation today.

Quoting Rheinwaldner (Reply 45):
I see another problem: The inside mandrel blocks also easy configuration changes that could arise from customer wishes (and in small numbers).

You don't change primary structure on a customer wish (unless the customer is willing to pay for it, and it would be a *big* payment).

Quoting Rheinwaldner (Reply 45):
Things like changing the number of windows (have blind windows) or the number of emergency exits and such things can not be delivered by the basic frame.

Since, as I learned earlier in this thread, the barrels are done whole and the windows cut later, they could just not cut the windows or doors...it would actually be easier (not counting the aforementioned structural issues).

Tom.
 
astuteman
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RE: Cfrp Panel Fuselages Superior (?)

Thu Sep 25, 2008 6:05 am

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 in a year....  Smile

Rgds
 
rheinwaldner
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RE: Cfrp Panel Fuselages Superior (?)

Thu Sep 25, 2008 7:37 am



Quoting Tdscanuck (Reply 47):
Quoting Rheinwaldner (Reply 45):
I see another problem: The inside mandrel blocks also easy configuration changes that could arise from customer wishes (and in small numbers).

You don't change primary structure on a customer wish (unless the customer is willing to pay for it, and it would be a *big* payment).

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 more than loading another program to the tape laying machine. From the CAD drawing board transfer data digitally to the production line and create fully automated tailor made structures, what a vision! A machine able to realize special wishes at the same cost as mass production!
The panels for the freighter version are created just by selecting another radio-button on the machine control screen.

Like a flash this future vision popped in my mind:
If everything is realized this way and a high integration degree could be achieved (at least automated production of shell, stringers, frames and window-frames) the following vision could become true:
Development of a software models for really different configurations. The customer may choose things like window size, number and pitch and the fuselage software models creates the shell accordingly. "Creates" means calculation of the required reinforcements and fit in the size and number of frames for the choosen window pitch. After that the tape lying machine is loaded with the program and the shell comes into existence fully automated. Each airline can match seat pitch with window pitch! Or match the window pitch with the requirements from your first class suites! Or to some degree choose the location and number of doors or emergency exits. In the future such things may be not more than "software" solutions and to a lesser degree maybe certification challenges.
(This idea is quite exposed to be pulled apart by realists. I am aware that it is not more than a long term vision. There are a thousand well founded questions to be answered before such a method could become reality).

Quoting Tdscanuck (Reply 47):
Quoting Rheinwaldner (Reply 45):
Things like changing the number of windows (have blind windows) or the number of emergency exits and such things can not be delivered by the basic frame.

Since, as I learned earlier in this thread, the barrels are done whole and the windows cut later, they could just not cut the windows or doors...it would actually be easier (not counting the aforementioned structural issues).

Is the shell itself made stronger around the windows? That would be possible with barrels easily. But it would be like hard-coded in the inside mandrel. No possibility to leave windows away (you can but only with the weight penalty of the unused enforcement). For the freighter (where you don't want windows and their enforcements at all) this again means new mandrels.
If the barrel just gets a uniform shell-thickness at the windows (something you probably aim for with inside mandrels and barrels) you loose some (weight- and production-) potential of automated reinforcements.

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