Composite wing spar design: Shear web/spar cap interface

Discussion in 'Aircraft Design / Aerodynamics / New Technology' started by Autodidact, Jan 9, 2010.

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  1. Jan 9, 2010 #1

    Autodidact

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    I've been thinking a little about composite wing spars and something that Billski said in another thread.
    I am not trying to say how anything actually works, I am just relating my observations and tentative opinions about some VERY crude experiments I've done.
    First....

    What are composites?

    I think when we say composites on this forum, we generally mean long strands of fiber encased in a "plastic" matrix. The fibers by themselves do a pretty good job of reacting to tension forces, but by themselves, are not good for compression loads. They will buckle under their own weight. When we bundle many of these fibers parallel together and encase them in a plastic "matrix", the matrix together with their close proximity to one another makes them able to resist buckling and we now have a structure capable of withstanding both tensile and compressive loads such as we would want in a spar "cap".

    But what about shear loads?

    Shear is not like loads normal to the cross-sectional area of the spar (normal=perpendicular in mathspeak, i.e., tension/compression as in the spar caps described above). Here, shear is longitudinal along the spar, to put it very simplistically, parallel to the tension/compression forces, but it is different. When you push a sanding block along a piece of wood, the force that resists you is "shear" force between the block and the wood. Think of the top half of the shear web pushing the bottom half of the web toward the wing tip and the bottom half pushing the top toward the root. My first intuitive guess about a strand orientation for a shear web "weave" was a bidirectional with the strands at 90° to one another and 45° to the spar caps. So I cut out a square of aluminum screen "cloth" so that the strands were at 45° to the edges of the square(see fig. "A"). Then I clamped opposite edges of the "cloth" and put the square under a shear load and it seemed to do a pretty good job of resisting the shear load. It would deform to the shape of fig. "B", but not without some force and wrinkling, just like a solid foil sheet would have done. So I formed the opinion that 45° to the spar cap was a reasonable strand orientation for the shear web.

    What about tension and compression loads?

    Then I compressed one edge of the "cloth" and stretched the opposite edge so that the square then became a trapezoid as in fig. "C". This required no effort and there was no buckling or wrinkling. I concluded that my shear web "weave" was not good at withstanding tension/compression forces.

    Next, I stretched it out.

    I formed the square into a long rectangle so that the strands were now at about 20° to 22° to the spar caps and tried the shear force again and it did seem to react better against the shear force. Also, if I clamped the short edges on the ends, one of the long edges seemed to offer more resistance to compression force, but not a great deal more. If I unclamped the short edges at the ends, then I could easily compress a long edge and the rectangle would easily form a trapezoid, but not without a great deal of vertical expansion (see fig. "C"). I then formed the opinion that a good strand orientation for the shear web might be similar to that shown in fig. "D".

    But what about what Billski said?

    Billski has indeed said in several different posts that there is a problem with shear web failure where it joins or is influenced by the spar caps. Or at least thats what I think Billski said. I assumed that it was obviously because the shear web fibers are at an angle to those of the spar cap and this would cause them to stretch or shorten a comparatively greater amount than the spar cap fibers. But this is not true, the opposite is what actually happens, see fig."E". But then I remembered that the only place in a beam cross section that is not subject to normal (tension/compression) stress is at the neutral axis. From the neutral axis, stress increases from zero to a maxim at the outer sufaces of the spar caps. The neutral axis is a line drawn crossways through the beam cross section about which the beam cross section will generally balance.

    This means the shear web is under combined stresses.

    So, I simulated combined shear and normal stresses on the shear web cloth, i.e., I returned it to a trapezoidal shape (fig. "C") and then applied a shear load to it. This time it no longer did such a good job of resisting the shear load and also, there was a region in the center of the trapeziod where the "cloth" seemed to be trying to twist into a spiral shape. I you look at fig. "C", you can see that all of these strange deformations cannot be good for the matrix!

    A good solution seemed to be...

    As per Billski's recomendation in his post in the "Upsilon" thread, a good solution seemed to be to extend the spar cap down into the shear web so that the cross-sectional area of the "spar cap extension" was greater in comparison to the amount of force it had to deal with than than the cross-sectional area of the original spar cap needed to to be. Greater cross-sectional area for a given amount of force means less stress and, since stress and strain are directly related, less strain, i.e.,less "movement" (stretching/shortening) of the spar caps and less torture for the shear web. At first, this might seem like a waste of spar material, but the only alternative is to increase the shear web material and as we've already seen, it is not suitable for reacting to normal stresses. So it would require more material and be heavier. At least thats what I think Billski was talking about. A very fanciful sketch of what I imagine a properly designed composite wing spar MIGHT look like is shown in fig."F", hopefully someone will critique it and set me straight; I have steeled myself for the news. Even if I have got anywhere close to understanding this, without an understanding of the mathematics involved, I'm still just a beagle chasing a Farrari!

    And mathematics is the rub here.

    Heres what Raymer says about compsite structural design: "To do it right requires tensor calculus equations and a pretty good computer program." Tensor calculus! This is the mathematics of Relativity, Quantum Physics, Finite Element Analysis, Computational Fluid Dynamics, etc. Another book! Here's what Wikipedia says: "For the anisotropic material, it requires the mathematics of a second order tensor and up to 21 material property constants. For the special case of orthogonal isotropy, there are three different material property constants for each of Young's Modulus, Shear Modulus, and Poisson's ratio - a total of 9 constants to describe the relationship between forces/moments and strains/curvatures."
    Billski, in his rather long and informative post in the "Upsilon" thread, has made it clear that he does not favor simplified methods, specifically, Martin Hollmann's methods, over the more rigorous methods covered in books which no doubt require proficiency in understanding tensors. Raymer does recomend Hollmann, but with a caveat: "A good technical introduction to composite materials is provided in [Composite Airframe Structures by M. Niu]. For homebuilders, Hollmann provides an excellent overview of composites." I am not qualified to judge Hollmann's methods. Even if I had read his book, I would not be qualified to do that until I would have mastered tensors and read the more rigorous treatments. I will say this, if Hollmann's methods can lead to a strong enough, safe enough, albeit somewhat heavier than necessary wing structure, then this is not the worst case scenario, for the non-degreed among us who have no tensor calculus, which would be no wing at all. And I can see a time coming when, in a last ditch effort to avoid learning tensor calculus, I will probably buy some of Hollmann's books. But I can also see that having read Hollmann's "overview", curiosity will probably get the best of me (it always does) and I will crack open my Tensor Calculus.
     

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  2. Jan 9, 2010 #2

    wsimpso1

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    This is worthy of a longer session than I can give right now. More after I get some parts under vacuum...

    Billski
     
  3. Jan 9, 2010 #3

    Norman

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    Just one observation from the peanut gallery. Since a fiber under tension will try to straiten out your hourglass cross section might result in the webbing delaminating from the core. The dynamic soaring crowd has been using a compost spar design consisting of caps separated by a compression insert wraped with a kevlar sock, its cross section looks like what you've got there except for the tapered caps and the webbing is vertical. This design removes the compression load from the webbing which alows it to be very thin.
     
  4. Jan 9, 2010 #4

    autoreply

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    Just some non-connected issues:

    I always visualized the simplest spar a block of foam, two caps glued to it and then wrapped in 45 degrees glass/carbon.

    As for composite design; if you want to do it correct it's very complicated. As the Boeing guys have proven even the most refined math is still not up to the job of a real-life composite structure, since their calculations were wrong by 30%.

    From my point of view, if you apply conventional mechanics of materials on a simple structure and you have a thorough understanding of the behavior of composites you should be fine using them if you apply a safety margin of 2 AND if you test the most important parts.

    After all, the spars we normally use and the torsion webs we normally use aren't that complicated structures and they have a limited number of failures modes. Where it gets really complicated are 3D joints and well, the deviations in fibre orientation, part precision and such are usually big enough to simply make it more beneficial to just a good safety factor and test.
    When looking at the essential construction of our aircraft, that's usually limited to a spar and torsion elements (fuselage, leading edge, flutter basics)) that can be analyzed by "acceptable" mathematics.

    The latter is another reason to keep my personal design simple, that reduces design and production time while it makes the number of unknown variables (strenth) much smaller.
     
  5. Jan 9, 2010 #5

    mz-

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    Is it mathematically proven that the right shear web fiber orientation is +-45 degrees? I can see that it could be: if you make it to a shallower angle, you get longer fibers per load and more mass. If you make it steeper, the forces per fiber are greater as the alignment is more perpendicular.
    But if there's a derivation of it somewhere, it might be interesting.
    Also, because the positive and negative loading is different (and possibly top and bottom spar cap size), if you wanted to save weight, you could have more/less fibers in the +45 vs the -45 direction...
    I don't know how much compression strength is needed and how important buckling is in a spar cap. Is that actually the size defining issue?
     
  6. Jan 9, 2010 #6

    wsimpso1

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    Big topic!

    First off, the screen analogy is called meshing analysis, and is one way taught to warm up students to the idea that the fibers carry the load. But let's remember a few other things.

    Composites are structural components made up of glue and string, and quite frankly, the string should be aligned with the loads for the most efficient structures. Your intuition about making the shear web at an angle is very good. If the spar had only bending loads, +/- 45 degrees would be ideal. Interestingly, this is also the ideal angle for elements under pure torsion too. Once you get the resin in there, the difference between +/-45 and 0/90 for shear is still there but the difference is now one of degree instead of being night-and-day. Back to webs, the spar carries only bending moments in a few places - usually there is some shear as well, which can drive a design to slightly different angles, but +/-45 is available with woven and with knitted cloths. Yeah, you could lay it up with UNI cloth at +/- whatever angle you want, but +/- 45 is usually close enough. It also lays up nicely, whereas 0/90 may not want to follow the curves as well in open layups. If you are vacuum bagging, resin infusing, or using prepregs, this is not much of an issue.



    Be aware that a designer of note named Orion has some sophisticated perspectives on using 0/90. I forget right now if it had to do with fiber reinforced plastics or plywood (natural fiber and resin composite with orthogonal orientation) or both. Search on his name and "shear web" for his input on that.

    The resin matrix is there to bind things together, which allows all kinds of magic above what a meshing analysis would predict... First, it does allow much high compressive strength, as you inferred. It is also the thing that allows load from one place to be transmitted across the parts. You need a fair amount of distance along each fiber to pick up a load from adjacent structures - you can not just glue fiber ends together as the fiber can be ten to a hundred times as strong as the adhesive. It also allows (as was aluded to above) 0/90 to have shear strength.


    If you imagine a unidirectional layup - all fibers running the same way - we can make some interesting insights. And this is realistic, as every little bundle of fibers making up a strand in a woven cloth fits this description. Axial strengths (both compressive and tensile) of the lamina will be limited by either axial strength of the fiber or by the ability of the resin to transfer the loads. All of the common laminating resins and fibers are compatible in this way, so first fiber failure is close to the min strength of the raw fiber stock. Since the apparent area of the lamina is greater than just the cross section of the fibers, you end up predicting strength by computing the failure strain of the laminate, the failure load, and then dividing by the apparent section under load. Typically the resin is no where near its limits and has much less stiffness too, so if you took the first fiber failure strength of the fibers and multiplied it by the resin fraction, you would have a fair estimate of lamina strength...

    Take the same lamina and put it in compression across the fibers, it will be stiff and strong indeed. Put it in tension, and the stiffness will be the same, but it will fail when the stress in the resin reaches the tensile strength of the resin. Quite low compared to the tensile strength of the fibers.... Non symmetric failure envelope.

    Now try to shear the lamina. One way, the shear is trying to slide layers of fibers relative to the next layer, and you are really testing the the shear strength of the resin, which can be kind of low. Across the fibers, you are trying to shear the fibers, and they dominate, which is a lot higher. Yeah, this is the same as your screen at 0/90, but the fibers are supported by the resin, so that they all must move together. As an example of this, imagine a pack of playing cards, flexes pretty nicely for shuffling, right? Now imagine you glued the cards together. It still looks like the pack of cards, but now it is hell for both strong and stiff. Same thing happens when you cure resin around a woven cloth... This where the screen analogy breaks down.


    Now every strand of fibers woven into cloth behaves this way, and we were just describing on-axis characteristics. Once you start weaving a cloth or laying something up, you are adding different angled lamina together, and it gets complicated...


    So elastic characteristics of a lamina (fibers in one direction)... We need Young's Modulus along the fibers, Young's Modulus across the fibers, Poisson's Ratio, and Shear Modulus. Hmm. It becomes a 6x6 matrix with four spaces open. Transform it with loads that are not aligned with the fibers, and it fills out the whole 3x3 matrix. Don't worry too much about the term tensor math. It is matrix algebra, and all of the terms for the transformation matrix are sines and cosines of the fiber angle. At 0/90 , the terms are all 0's and 1's, while at 45 degrees, it all breaks down to 1/4's and 1/2's.


    Then comes the fun. Combining the lamina into laminates, the ABBD stiffness matrix, the NM load vector, solving for the ek deflection matrix.


    Now strength gets interesting. You start with tensile and compressive strengths along the fibers and across the fibers, and shear strength. In most materials, the tensile and compressive strengths along the fibers are the same, but some materials (Kevlar is one) compression strength is much lower. And almost all materials have much more compressive strength across the fibers than tensile strength across the fibers, and both of these are much lower than with the fibers. So, rather than have a nice oval failure envelope like metals and plastics, we have a very flat oval that is also asymmetric. This gets transformed into six terms for strength that can e applied to the deflection. In some directions, you material can not take anywhere near as much load as it can in other directions...


    Whew! That is a whole senior/graduate level college semester course and they assume that you start with the matrix algebra already in your head... When you start using the gross simplications of Martin Hollmann and Andrew Marshall, you can find that your structure is severely underbuilt when you get to a place in the process where you can test things. The 30% errors at Boeing were in complicated structures and could be fixed pretty easily – Hollmann designed wing spars for Lancair were weak by something around a factor of two, and they ultimately fixed it by tripling the shear web... You can not test everything, so if something ends up understrength, you had better have enough altitude to get out and be decelerated by the parachute. Good analysis is a better scheme, and good analysis is more trouble to do.


    Now, back to your original question... In designing a composite spar, we will build a spar with the required bending stiffness and strength at less weight if we make the caps relatively thin and wide. This maximises the distance between caps, which allows minimizing the cap cross sectional area. Next, one method is to fill the space between caps with plastic foam, and then wrap shear webs on both sides and onto both caps. This spreads the shear load into the cap over a large area and the foam stabilizes the webs against buckling. You have to check that you are within the bond strength of the resins you are using, but this rarely drives changes. The other method is to build either a channel or an I beam. Either way, a foam shear web is installed and the shear web is laid up bridging between the caps and laps onto the inside faces of the caps, again spreading the shear load onto the cap over a considerable area. In the case of the I beam, you do this with the shear webs centered. In the case of the channel, the web is shoved to one side, and a couple plies will be taken around the outside of the caps and onto the outside of the shear web too. This ties everything together, and provides further support of the shear web against buckling. And the big thing to make sure of is that the shear load between the cap and web is spread over enough area to be below the strength of the resin serving as adhesive...


    Others things to check are that your interlaminar shear is always within the capabilities of the resins present. When you build a cap like in your illustration, you start to approach issues of shear strength of the resin between caps and webs, and between lamina in the caps. In truth though, the biggest drawbacks to that design is it will be heavier than with wide thin caps and it will not bond as nicely to the skins with the narrow contact.


    Now in building wing spars, you can probably start with Martin Hollmann's approach, but you need to remember a few things and you need to at least learn how to compute the wing axial and bending deflections and how to check failure criteria:


    First, the wing spar sumps the wing shear into the fuselage wall, not at the centerline. Determine your shear and bending diagrams with that in mind;


    Next, do the rough sizing using Hollmann's methods;


    Compute spar stiffness and loads and from that axial and bending strains for every laminate at both caps, and at the top, middle, and bottom of the shear web;


    Check the failure criteria of your materials for every lamina at all of those places where you have the strains.


    Once you do this, you WILL find that the shear webs are understrength. To have a chance at this stuff, I set it all up in Excel, and then do one station along the wing at a time. I put in a space for spar depth, cap plies, web plies, and all of the beam computations, ABBD matrix Gaussian matrix solution and failure checks. This allows you to iterate the numbers of plies of each and find the combination with minimum cross sectional area that does not fail any part. You will add plies to both the caps and the web and repeat the calcs until the failure criteria is satisfied, and then record the plies of each for that station. Do this over stations every 6 to 12 inches over the half span, and you have your lamination schedule for caps and webs.


    Ugh. Just using Hollmann's methods straight will result in understrength parts. And you will not know how much to beef them up. While that method might make design limits on a test, you won't have the intended factor of safety, which is where durability and robustness over time really comes from...


    Remember too that aero loads will come in from the skins and ribs and attached to the spar in less than perfectly distributed ways, and those margins above basic strength are in place to help with that too. A factor of safety of 2.0 is in composites for good reasons...


    So, chew on that for a bit, and see what you think. We can certainly spend more time on this conversation. By the way, there is no calculas in using the method, but you do need some numbers for your materials and matrix algebra.



    Billski
     
    karoliina.t.salminen likes this.
  7. Jan 10, 2010 #7

    Autodidact

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    Oy! Thanks, Billski. I had to compose and write my post twice because I think I took so long the first time that my log in expired. Anyway, I lost my first post and had to write it all again. Rage. Sadness. Denial. Acceptance. That took a while and then I wrote it all over again. So believe me when I say that I really appreciate your taking the time to write these dissertations. I will be thinking about this for quite some time, so don't be surprised if I don't say anything about the subject for a while.
    One thing, though - could you post a good picture or sketch that illustrates one or two correct examples of a cross section of a composite spar? And I see the problem that you and Norman pointed out with the web wanting to delaminate from the caps I drew as well as the caps being too heavy and being better when they are wider and thinner. It is difficult to get the homogeneous material paradigm out of my head, and I don't even know that one terribly well!

    Edit: Actually, I just realized that I keep trying to visualize a WOODEN spar instead of a metal one when I think of what to compare composites to. Metal spar caps tend to be wide and thin too.
     
    Last edited: Jan 10, 2010
  8. Jan 10, 2010 #8

    Autodidact

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    This is all making me realize how very little I understand, on the other hand, I'm becoming more and more intrigued by this type of structure. It looks like 45° is the best compromise between angle and number of fibers actually connecting between the caps.

    Edit: Oh, wait - the matrix constrains the fibers and shears them over their cross sections, ±45° offers a greater cross-sectional area for shear resistance. Like Billski said, the screen analogy breaks down.

    I believe that since a column under compression will buckle long before it reaches yield strength of the material, buckling load is the defining issue. On the other hand, a spar cap in a wing is supported by the web, ribs, skin, etc., so I think that may be only partially right. Engineers like to err on the side of safety so it might be considered the defining issue just to simplify things.
     
    Last edited: Jan 10, 2010
  9. Jan 10, 2010 #9

    Autodidact

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    Since a foam core can provide acceptable compression strength, and since FRP is complex for shear strength, I wonder if there is a homogeneous material that will bond nicely to FRP spar caps? Like, say, ABS plastic? Ah, but it needs to bend so that it has a large area to bond to those thin, wide caps. Hmmm.
     
  10. Jan 10, 2010 #10

    GarandOwner

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    This is not always true, sometimes a spar may fail due to buckling, but sometimes it will fail at its yield strength due to the bending stress. So both cases must be examined when designing and testing your design. You cant just assume that it will fail due to buckling and only test/design for that mode of failure.
     
  11. Jan 10, 2010 #11

    wsimpso1

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    Lots of stuff...

    Relatively wide thin caps make for a more efficient spar (It gets to strength at lower weight). You do get into diminishing returns as the cap gets thinner, and then as it gets wider and thinner, you can get into cap buckling along the free edge. Jones talks about that a bit in his book, and Bruhn has quite a discussion, but you would have to do some translation as he was discussing aluminum...

    In most wing structures, the spar is bonded to the skin, and then there are usually some ribs, although they are pretty far apart in composite wings. Anyway, the skin and the ribs all tend to support the caps against buckling, so the buckling mode that is left is between ribs and the skin has to buckle with the spar cap. In little fiberglass airplanes with open structure spar and rib type wing structures, 3-4" wide spars are about it. In box beams (Long EZ and derivatives center section spar), they go wider, but then they have a web on each side of the cap, so the buckling mode becomes even harder to get to. In foam core wings (Long EZ and derivatives) the channel spar is fully supported with core foam and the skin too. If you stick with monkey-see monkey-do design methods on design and proportions, you might be OK, but otherwise I would spend some quality time with Bruhn here to figure out how to avoid cap edge buckling.

    Homeogeniety paradigm is one that you need to get out of, although once you do, it frees you up to think about all sorts of neat stuff. Some of which might be really hard to actually build...

    There is an analytical tool taught to mechnaical engineering students in the sophomore year call Mohr's circle, and it is a way of representing the stress in an element including the orientation and allows the calculation principal stresses. Using it to examine pure shear elements and pure torsion elements, it has been shown that +/-45 degrees is the angle of principle stresses, and that is thus the ideal angle to orient fibers for pure shear. So, yeah, it has been proven. And every BSME in the world has been exposed to it atleast twice, once in Mechanics of Materials, and once in Machine Design.

    Billski
     
  12. Jan 10, 2010 #12

    AVI

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    Billski,
    Are you familiar with Jim Marske's Composite Design Manual?
    Any comments regarding his methods of spar design calculations and construction
    methods using Graphlite carbon rods?
    Alex
     
  13. Jan 10, 2010 #13

    wsimpso1

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    I am familiar with Jim Marske's sailplanes and his website. I am sure that with graphlite rod, an amatuer can take advantage of carbon's strength in spar caps. I have not read his book, so I can not comment on his methods.

    Can anyone tell us what Jim Marske's book has to offer? How does he size spar caps and shear webs?

    Bill
     
  14. Jan 10, 2010 #14

    Mac790

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  15. Jan 10, 2010 #15

    wsimpso1

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    I just went back in and looked at the posted pages from Marske's book. Same technique as everyone else uses for spar cap sizing. Figure out the bending moment and depth of the wing, put in a max stress in the cap, get out a cap area.

    Anyone know if he was similarly conventional for sizing shear webs?

    Bill
     
  16. Jan 12, 2010 #16

    Michealvalentinsmith

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    Here's some of what I saved from the web page. Seems it all been removed now.

    I had hoped to upload it all for you here but it looks to be too much and the jpgs are no the right size for this web site - and I'm not so keen as to want to resize them manually.
     

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  17. Jan 12, 2010 #17

    Michealvalentinsmith

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    Regarding amateurs Marske claimed he had inexperience women build the spar is the aluminum form faster than experienced men.

    I've attached a file where on guys added some graphlite rod to a wooden spar in a simple key slot. The G load went from 3 to 7.


    Carbon rod alos available from :

    Avia Sport Composites Inc,

    Introducing a new construction material system to build composite aircraft wing spars. Whether your spars are constructed of fiberglass or carbon rovings you must take a serious look at this new miracle structural material. It's more than six times stronger than 2024-T3 aluminum, twice as stiff and nearly half it's weight. Compare this remarkable premanufactured carbon rod with wet layup carbon roving, it's nearly 3-1/2 times stronger in tension and 5-1/2 times as strong in compression.
    It's the most exciting material to hit the aerospace market since the introduction of fiberglass. Under the trade name of GRAPHLITE this material is cost effective and consistant in properties. You can build wing spars faster with much greater reliability. Fabricatedt in one operation in a female mold, it cuts assembly time in half. Another benefit of a female mold is that the outside dimensions are always consistant.
    The current problem with wet hand laid up carbon roving is that you cannot lay all the filaments down straight or achieve a consistant resin content. As a result your laminate strength has a lot of localized strength deficiencies.
    Here is how this problem was solved. A new highly modified pultrusion process, forms a round or rectangular carbon rod in a machine which lays in all filaments straight, parallel and under equal tension. Resin content is closely controlled to +/- one percent. Maximum performance is obtained in every fiber resulting in tensile strengths exceeding 320,000 psi and 275,000 psi compressive strengths, far above the nearest contender. Coupon testing has shown consistant strength values with very small scatter. Modulus of elasticity is 21.5 million.
    The GRAPHLITE carbon rods are rolled off a spool, cut to length and laid into a female wing spar mold and embeded in an unfilled epoxy matrix. The rod ends are cut off square with a dremel cut off wheel or a fine tooth hack saw. The rods do not require cleaning nor sanding to improve bond strength. Rather than weaving glass fabric through the rod pack we let the epoxy matrix carry all the shear loads. This is similar to steel reinforcing rod set in concrete for structural beams or roadways.

    After working with 'GRAPHLITE' carbon rod for some five years now, I am still impressed with it's incredible strength and light weight. I compared prices of comparable spar cap materials of aluminum, wood andcarbon roving to see what the price difference might be. Would you believe that the carbon rod material is half the cost of all other materials? You see, so little GRAPHLITE is used that the cost is not significant. To give you an example. The new Monarch G, a 42.6 ft span sailplane, uses only four rods (.092 x .220) in each spar cap. Enough rod for one wing panel (upper and lower caps) requires 112 ft. That amounts to 1.44 lbs of rod per wing spar. The total weight of the completed wing spar is 7 lbs which includes blocking for root and strut fittings.
    Mat Kollman worked with us on the Monarch project. Mat is a hang glider pilot and recently designed a rigid wing hang glider of 42 ft wingspan and a cantilevered swept back wing called the Raptor. It utilized the same spar and D-tube as the Monarch G. Mat static loaded his cantilevered 'Raptor' wing to a negative 5g without failure. The stress at the root on each spar cap was 155,000 psi at 5g. This is half the stress of what this material is capable of carrying.
    Because the Monarch uses a wing strut the bending stresses on the wing are not as great as the Raptor's. The strut intersects the wing 5.5 ft out from the fuselage. Calculating the rod stress at the strut junction at 155,000 psi comes to 10.2g. Despite all this extra strength the spar cap weight in either of these craft weighs less than 1.5 lbs per wing panel. Samples of the rod have been tested at a laboratory to 326,000 to 333,000 psi in tension. Compression failure excedes 275,000 psi. Lab tests on hand laid up carbon rovings failed in the range of 47,000 psi to 80,000 psi in compression. A shocking realization.
    The Genesis was the first sailplane to use the 'GRAPHLITE' carbon rod. The Genesis is a 15 meter (49.2 ft) production sailplane with a slightly swept forward wing. During static testing a completed wing the test holding fixture failed at 19g. Test conditions were: a 260 lb pilot and 50 lbs of equipment. Extensive dynamic cyclic testing was also performed. 20,000 cycles in positive and negative loading, many above the design limit, failed to show any visual sign of rod failure or separation from the binding matrix.
    The prototype Genesis sailplane had a hand lay up spar from contineous carbon roving. The spar weight was 38 lbs. The production Genesis with the GRAPHLITE rod spar weighed only 25 lbs but was twice as strong. A third generation spar of GRAPHLITE for the Genesis would weigh under 20 lbs and be of equal strength. Just for the record, an equivilent aluminum spar would would weigh 65 lbs and be half as strong.
    The Lithuanian engineers were so impressed witnessing the Genesis static and dynamic test that their new sailplane, the 18 meter (59 ft span) LAK-17A, incorporates these rods in the mainspar also. What impresses me is the spar depth at the root of the LAK-17A is only 3.5 inches. Aspect ratio is 33 to 1. And the wing panel weight is only 122 lbs.!!
    The spar, such as used on the Monarch G and Raptor, is easy to make - even for a one-off aircraft. Bend up a piece of metal sheet with 1.5" high flanges top and bottom. Use this channel as your mold. Wax and PVA the mold. Lay in the channel one layer of 9 oz bid cloth, with weave at 45 degrees, and brush into it a coat of epoxy. Next place a carbon rod, full length, into each corner and laying flat. Wet out the rod well with epoxy resin. Lay the second rod on top of the first and wet out. This rod is shorter. Lay in the third rod which is shorter still. Add a fourth and perhaps a fifth depending on strength requirements. Now lay in a layer of 8.6 oz bid glass cloth into the channel and up the flanges packing the cloth tightly around the two stacks of carbon rods. It is recommended that the rod pack always be completely wrapped in glass cloth to prevent delamination from the shear web.

    If you wish, two channels may be bonded back to back to form an 'I' beam for heavy duty spars. A 6mm PVC foam sheet can be bonded between these two channel spars to act as a vertical stiffner.
    The stickler here is deflection. Because so little carbon rod is required and the material is so strong the beam just keeps bending as the load is increased. Don't let the high modulus fool you, the modulus of elasticity of GRAPHLITE (21 mil) is twice that of aluminum (10 mil ). We can illustrate the deflection if you look at the material stretch in tension at yield strength. Aluminum (6061-T6) has a modulus of 10 mil and a yield strength of 35,000 psi. Divide the modulus into the yield strength and you have (35,000 / 10,000,000 ) .0035" material stretch per inch of length at yield. For GRAPHLITE it's (320,000 / 21,000,000 = ) .015". Thats ( .015 / .0035 = ) 4.29 times more deflection for GRAPHLITE over aluminum. For sitka spruce it's: (9,500 / 1,300,000 = ) .00731". Here GRAPHLITE is 2.05 times as flexible in tension. For unidirectional fiberglass rovings it's (80,000 / 5,000,000 = ) .016". And thats nearly the same as GRAPHLITE. The difference here is; You are using 4 times less material (in carbon) plus carbon is 20% lighter per cubic inch than fiberglass. So you have, by comparison, for the same strength and deflection, a cap weight of 20 lbs in glass but only 4.2 lbs in GRAPHLITE.
    If you are building a short thick wing the deflection is probably not a problem. However, sailplane wings are long and thin. By doubling the number of rods in the cap the deflection is cut in half. You can do this by degrading the rod strength in half to 160,000 psi. However, you are still way ahead in weight and cost savings.
    Some builders want to mix the rod into a wood spar. Yes, you can do this. It's not as weight efficient as an all carbon spar but the little weight added is far offset by the strength gained. How much rod needs to be added? Lets say you have an I-beam spar with an upper cap cross sectional area of 1 square inch it would take .05 square inch of GRAPHLITE, or 4 rods of .125" dia, to double the strength of the spar.
    You need not scuff the surface of the rod prior to bonding. In fact, no surface preparation is required. Most pultruded rod has oil added to the resin to assure smooth flow through the dies. The GRAPHLITE manufacturing process does not require any foreign matter to be mixed into the die during the cure cycle. The matrix, or binder, of the carbon fibers is, Bis F Epoxy. Remember, epoxy likes to bond itself to epoxy.
    GRAPHLITE carbon rod is stronger than all other carbon because all fibers run in the same direction, are perfectly straight, and are under equal tension. Since it is machine made the resin content is closely controlled and laboratory tested breaking strength is repeatable with very small variation.
    The GRAPHLITE carbon rod comes in many sizes but the two recommended sizes are .125" (3.16mm) diameter and .092 x .220" (2.34 x 5.59mm) rectangular rod. Cost of the GRAPHLITE rod varies and depends on the quantity purchased. Small quantities, under 100 ft, may run $1.00 per foot and quantities of 500 ft, is $.65 per ft.
    Where do you get this wonder material? There is only one factory in the US that makes GRAPHLITE using this special process. It's inventor, Reggie Durhman, says it's a highly modified pulltrusion process. So if you want to try this great stuff get in touch with me at: Marske Aircraft Corp., 975 Loire Valley Drive, Marion, Ohio 43302. Phone: (740)389-3776, FAX (740)389-2236, or e-mail <marske@marion.net
     

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  18. Jan 12, 2010 #18

    Lucrum

    Lucrum

    Lucrum

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  19. Jan 12, 2010 #19

    berridos

    berridos

    berridos

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    What about the rear spar?
    I couldn`t find any info on the requirements and building technique of it.
    Should it also be 45º to absorb the torsion, with the same sandwich laminate thickness as the wing skin?
    Due to the corners, reinforcements at the edges to the wingskin?
    Reinforcements at the holes for the hinges?
     
  20. Jan 12, 2010 #20

    berridos

    berridos

    berridos

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    I haven`t seen anybody using nomex honeycomb in the main wingspar. Why that?
    I worked in marine applications with it and it is really amazing. Nothing to do with sandwich foams.
     

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