Hello everyone, it’s been a while since I’ve posted anything on the forum, although I have been popping in once in a while for a quick read of the interesting things going on. I’ve been meaning to write this post for some time now as it’s quite an important subject and I haven’t seen it discussed too much here. In the last few years, I’ve been working more heavily with the design of composite structures, primarily for certified aircraft (development of STC modifications). As such, a full substantiation of structural strength of the new parts/structures was required to the satisfaction of the aircraft certification authorities. Design was done using analysis and testing was required for the official substantiation since, as most of you know, the material properties of composites are highly dependent on the combination of the materials, manufacturing processes and environmental conditions. Testing of composite structures is not as straight forward as testing traditional metal structures. Metals have well known properties over a wide range of operating temperatures. The published properties databases (such as the FAA MMPDS document or its predecessor MIL-HDBK-5) have statistical backup such that the values can be applied to say for example any 2024-T3 sheet or 6061-T6 extrusion etc. Therefore, the metal structure can be tested in the shop under typical environmental conditions (unless the design calls for high/really low temperature applications), subject to the limit or ultimate loads. Ultimate loads are simply the limit loads multiplied by a Safety Factor of 1.5 as specified by FAR 23.303, and any other Special Factors (1.15 for fittings etc) as specified by FAR 23.618. Easy. With composites, it is required that the effects of variability in manufacturing are taken into account as well as environmental conditions such as hot/cold and dry/wet. So how do you do that? Some use a catch-all Safety Factor of 2 as opposed to 1.5 between limit and ultimate loads. I had come across it in the past in an FAA Advisory Circular (AC107?), but have not been able to find it again because I believe it has since been removed and is not accepted by the aircraft certification authorities. After doing some research on this matter, which I discuss below, I proved to myself why. (Note: this an example only) Material properties databases available to the public are quite limited, and mainly consist of materials such as pre-pregs, and process involving autoclaves or elevated temperature cures. None-the-less, a typical data set will provide values of strength, such as tensile strength, at varying temperatures and moisture saturation levels (wet/dry). From that data, a material performance envelope can be graphed. The one in the picture below is made from data I was fortunate to find in a military report for wet-layup room temperature cure 7781 fiberglass/Hysol EA9396 epoxy laminate. The points of the blue shape represent the ultimate tensile strength under Cold/Dry, Cold/Wet, Room Temp (RT)/Dry, RT/Wet, Hot/Dry and Hot/Wet conditions. So it’s clear that the tensile strength of the laminate can vary significantly (cold/dry strength is 5 times greater than hot/wet). It might be tempting for someone to look at the tabulated data and pick out the RT/Dry tensile strength value and apply it to the analysis of their design, but as you see, that’s not representative of the conditions an aircraft will see in service over its life. To establish proper design values, the design operating temperature range must be defined as well as an anticipated worst case saturation level. Fully saturated is overly conservative, and according to Michael Nui (Airframe Structural Design) a typical value of 67% saturation should be used. This is shown by the red line in the material performance envelope below. The temperature range to use is up to the designer. Remember the aircraft can be sitting and baking in the sun for a while. For example, if the maximum allowable temperature is set to be 60 deg C (maybe a bit high), then the design tensile strength at 67% saturation would be approximately 26 ksi… that’s far from the RT/Dry value of 53 ksi (knock-down factor of 53/26 = 2.04). For the case above, the ultimate tensile of strength of 26 ksi should be used for analysis purposes. If a load test is to be performed at room temperature conditions, with an unconditioned “dry” test specimen, the applied load(s) would have to be scaled by an overload factor of 2.04 (as calculated above) to account for the degraded strength at the worst case design temperature/saturation level. Therefore the ultimate test load would have a total Safety Factor of 1.5 x 2 = 3.0 over the limit load. This is just an example, as not all composite layups perform as “badly” at worst case conditions. From the data I have studied, it was clear that wet-layup, room temperature cured laminates typically have lower maximum strength and the most degradation in adverse environmental conditions (with more advanced materials/methods, the previously mentioned total Safety Factor of 2 is more applicable). What’s important is that these “low-tech” methods traditionally used to build many composite homebuilts. So if you are designing your own, careful consideration of the design safety factor you use is warranted, and should be representative of the materials and processes employed. In lieu of test data, some aircraft certification authorities have accepted Safety Factors of 2.25 for non-structural components and 2.5 for structural components. There is no reference for these values, but these may only be acceptable for one-off type modifications, not production products. Some of you may disagree with the method I’ve shown here to compensate the safety factor for composite materials, but the “moral of the story” is that I want to make amateur designers aware of the necessity to use higher Safety Factors when designing for composites, and to generate discussion on this matter. Maybe others can share their methods and how they approach this problem. If you made it this far, thanks for bearing the long post.