Looking at the cited article and translating for both those using SI and British units:
- 40 daN/mm^2 = 400N/mm^2 = 400MPa = 58.1 kpsi (carbon composite);
- 25 daN/mm^2 = 250N/mm^2 = 250MPa = 36.3 kpsi (glass composite).
The article stated clearly that stress levels below these with care to eliminate stress concentration is considered adequate. These are thus considered by whoever it was that wrote this item to be fatigue strengths, which are inhereltly much lower than whatever you want to call your stresses at single cycle failure. The article also goes on to tell us that higher stresses necessitate fatigue tests and or fatigue analysis to use higher stresses. Their guidance seems pretty clear to me.
Now this engineer finds this guidance troublesome from a couple directions:
- Unidirectional E-glass in my testing has a a compressive failure around 80 kpsi and the literature I have found cites safe minimum tensile/compressive strength at 70 kpsi. A safe fatigue strength for this same unidirectional E-glass might indeed be around one-half of ultimate strength (metals Sf start at 1/2*Su and work down), or close to the 36 kpsi stated. But if instead, you have a bidirectional laminate of the same material, strength measured on a coupon will be much lower and accordingly allowable stresses for good fatigue life will also be much lower;
- S-glass has both higher strengths and moduli than E-glass and supposedly higher allowables for simple design. So is the guidance provided for E-glass or S-glass? Both? For what kind of lamination schedules? It simply does not say;
- Unidrectional graphite structures claim much higher strength than in E-glass. Pre-pregs laminated and cured in an autoclave and pultruded rod (Graphlite, etc) claim strengths in the 200-300 kpsi range. Graphite laminates are also known to have very good fatigue resistence, so I would imagine we could design unidirectional graphite laminates to operate safely at much higher stresses than the 58 kpsi cited in the article. But what of bidirectional laminates? What of open layups, vacuum bagged or vacuum infused? These generally have much lower strengths and correspondingly lower fatigue strengths.
It appears that the guidance cited is perhaps not conservative enough for glass laminates and is very conservative for graphite. For this engineer's use, that information is just not useful... We would need to do more than simply compute stresses and say if we are good or not.
As to buckling etc. The classic airplane design texts cover buckling for aluminum very well. Composites are different. In composites, we have high modulus face sheets and cores and go for generous margins on buckling. I have gone with the stress and deflection formulae from Roarks, take the form apart to find EI and thickness, then use EI and thickness for composite panels to compute stresses and deflections in the panels. These formulae are for flat panels and we use curved ones, so they are again more conservative. In my various skins/cores, deflections get downright tiny, stresses modest, and buckling appears firmly at bay using elastic stability models. On top of that, my structures pass the monkey-see monkey-do tests of comparison with other airplanes of similar performance and good history.
Billski
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