Ply Orientation in Composite Laminates

bdk suggested this topic, and I also like to know more about it. Despite all the difficulties with composite, there is no contest of appearance (and aerodynamic drag) between a metal airplane and a composite airplane. So we have to deal with composite.

The textbooks on composite place a lot of rules on the "right" ply orientations and sequences. For example, you should have symmetry so the structure will not warp under tension. Another rule is that you should not have only uni-direction plies of the same orientation, otherwise the laminate has poor load bearing properties in the transverse direction.

I have seen both rules violated in spar and skin of wings of small composite airplanes. In one airplane, the skin has a glass ply, a carbon ply, a foam layer, and a glass ply (from the outer to inner skin), there is no symmetry in the laminate. The spar caps are solely made of uni-directional carbon of the same orientation, sometimes up to 20+ plies.

Are these ply sequences typical for small airplanes? I would like to know more examples of wing skin, spar cap and fuselage skin ply sequences. What are the "penalties" for breaking textbook rules?

The answer lies in understanding the stress on each member and orienting the weave to the stress. Some locations have most of the stress in one direction and little in the cross direction. Some locations have stresses in multiple directions.

The homebuilder generally can purchase cloth that is either plain or crows foot weave, or tape. The magic is figuring out where the stress goes and matching the direction of the fibers to that direction. And of course figuring out strength in both compression and tension. Even tapes have some cross weave to hold it together. That can be good enough when embedded in epoxy resin.

May I suggest that you purchase a copy of Burt Rutan's original text - Moldless Composite Construction. It is a wealth of practical information for the novice.

Commercial airplanes are build completely differently than homebuilts. Boeing and its peer companies invest in machines that can lay down almost individual fibers into a mold that then has the mixed resin injected, vacuum and temperature applied, and a while later a minimum weight, maximum strength assembly pops out. The engineers spend a lot of time figuring out exactly how much of what, at what orientation, produces the optimum combination of weight, strength, damage resistance, etc. You and I work with much grosser materials, techniques, and design safety factors.

Get Rutan's book and build a piece of a wing. Or sign up for one of the SportAir workshops. It can be rocket science, but it does not have to be.

Best of luck,

Wes

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Originally Posted by wantobe

What are the "penalties" for breaking textbook rules?

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What textbook are you referencing?

[QUOTE martymayes

What textbook are you referencing?[/QUOTE]

I do not have Niu's book at hand. I just searched "10% rule laminate" on the web, and found a paper with the following paragraph on laminate design rules. The first rule requires 0, +-45, 90 degree plies all be used and have at least 10%. I do not think we can afford it in fuselage or wing skin. The second rule prohibits the way the wing spar is constructed in small GA airplane, quite often with more than 20 uni plies. The third rule deals with impact damage. Again it is massively violated in homebuilt practice.


Laminate design rules

The bi-level optimization has been applied to include
laminate design rules (Niu 1992). Three such rules, which
are subsequently referred to as Rules (1–3), are as follows:
1. The thickness percentage of each orientation is controlled
with respect to the total laminate thickness. For
example, it is required that each laminate contains at
least 10% of each orientation. This constraint ensures
that laminates have sufficient damage tolerance, aeroelastic
stiffness, bolted joint strength, and the capability
to carry secondary loads. This rule is simplified by
giving percentage targets for each orientation. For
example, the target for the 0°, ±45°, and 90° materials
in the skin laminate might be 44, 44, and 12% and in
the stiffener laminate might be 60, 30, and 10%,
respectively.
2. The maximum number of successive plies in any
orientation is limited. For example, it may be decided
that four consecutive ply layers with the same orientation
is acceptable, but five are not. This constraint is
applied to reduce transverse shear stress and minimize
edge splitting.
3. The outer plies of the skin and stiffener laminates are
forced to be ±45° plies. This constraint is to improve
damage tolerance after impact.

I do not read that doc the way that you do. I read it as stating that for a specific type of laminate used in whatever the designers application is, the engineer must identify rules at the start of the design process that allow you to stay within the mechanical constraints (strength, fatigue resistenace, damage resistance, etc) of the materials that you are using.

I read general guidance, not specific guidance. The limitations on using kevlar vs carbon fiber are different. The limitations for West System epoxy are somewhat different than Aeropoxy. So when you start a design, you identify the strengths and weaknesses of the combinations of the materials as establish rules to use for applying those materials to your design.

Hope this helps,

Wes
N78PS

For most structural applications like fuselage, you want a pseudo-isotropic laminate of 0/90, +/-45, 0/90 for woven fabric. This orientation gives the most uniform strength in all directions and approximates a metal with isotropic properties. For specific beams like a spar, you want uni fibers in the caps at 4 ply increments and then a +/-45 ply to provide shear stability and to keep the uni (0 deg plies) from developing transverse cracks due to resin shrinkage, if part is thermally cured. If all layup is room temp cure, this is not as essential. For wing skin structure which sees both bending and torsion, you want skins to have equal number of 0/90 and +/-45 plies, the 0/90 for bending stiffness and the 45s for torsional shear strength. Usually, the 45s, (called bias plies) are interior plies and 0/90s are designed as surface piles. You worry about balanced plies ( symmetrical about centerline or neutral axis) when you have a thermal cure and an unbalanced laminate would warp; not an issue with rt cure systems.

What are the 'rules' aimed at? Large aircraft or smaller GA types? Probably large aircraft as those manufacturers can afford to research into these issues. Take a look at glider spar caps - often uni-directional with a wrap (+- 45 I think) to provide torsional stiffness. If you can afford the weight to have many piles in the skin then the 10% rule is valid. If only 2 plies are required each side of the core then something has to give. If the airplane you looked at is successful with several hundred examples flying then the designer probably made a sensible compromise. At least 10% in any direction implies at least 10 plies - many light aircraft don't have that many plies in wing or fuselage skins. What does Burt Rutan say?

Peter

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there is no contest of appearance (and aerodynamic drag) between a metal airplane and a composite airplane. So we have to deal with composite.

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I'll give you the drag one but some of us really happen to enjoy the appearance of a well-designed riveted metal or fabric aircraft. Then there are a lot of metal or fabric designs that are G-d awful. A lot of composite aircraft look like a sperm with wings attached while others are things of almost unrivaled beauty. When it comes down to something purely aesthetic in nature, there will almost always be a contest. ;)

BTW, do you have a copy of Niu's book? I have a spare copy around here somewhere that I picked up along the way if you'd like it.

The number of plies depends on the required strength of laminate in a probable loading/stress direction. You don't add plies if a single ply will suffice. For sandwich construction, a single woven ply on each surface of a core will carry both bending and shear stresses, if either is not too high. My Pulsar has one-ply skins over nomex honeycomb core and provides a very efficient/stiff structure. The weight of structure per sq ft is .258 lb with the whole plane's structural weight at <450 lbs.; with engine and avionics, weight is 630 lbs.
The 10% rule mentioned earlier pertains to the contribution of each ply NOT in the direction of the weave. A 45 deg ply would make only a 10% contribution to a 0 deg directional loading but a 100% contribution in the 45 deg direction.
This forum may not be interested in technical details of composites and having worked in the field for 40+ yrs in manufacturing composite structures, it's hard not to comment when you read questions about the technology.

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Originally Posted by wantobe

Despite all the difficulties with composite, there is no contest of appearance (and aerodynamic drag) between a metal airplane and a composite airplane. So we have to deal with composite.

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Not sure that is an accepted or accurate premise. Bede and Wittman proved you can push a square box through the air pretty damn fast if it's designed correctly and it doesn't seem to make much difference if it's aluminum or steel, wood and fabric. Dennis Polen built an aluminum airplane that has never been outrun by a composite airplane in the same class. Dr. Brokaw built an aluminum airplane that was so fast it was named "Bullet"...lol.