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This Excel spreadsheet is used in preliminary design on triple tapered vacuum bagged wings for weight and strength estimation. There are a few assumptions to be aware of. First, it assumes that you will be using a good quality foam such as blue foam or Spyder foam, so as to reduce buckling failures of the skin in compression. Second, it assumes that you are using material for the full chord, no triangular darts of material. Third, it assumes that you will be careful in your epoxy application, no extremes, just that you are not pouring on the epoxy. Fourth, it assumes that all fiberglass is put on in equal amounts, top and bottom, but allows different lay-up schedules for the top and bottom for carbon fiber. What is not included in the weights, is ribs for joiner load distribution, tubes for joiners, etc.
All input data is in yellow.
All output results are in "pink" (kinda...)
The units are typical english units, inches, oz/yard^2, ft-lbs, etc.
The usual way to run the spreadsheet is to modify the wing planform to match your proposed new creation. It has to havethree panels or less. If you are using a two panel wing, break the inner panel into two panels, and if one panel, break it into three panels. Make sure that you modify the wing planform shape factor if you are using something other than the typical efficient 3 taper wing that people have been using lately.
Then, put in the airfoil characteristics, the thickness, and the form factor. Almost all airfoils of interest have form factors from .6 to .65, so leaving this defaulted is typically Okay.
After that, you need to get the maximum load that the wing will see into the program. There are some inputs on Cl Max, Max airspeed, and max line tension in the upper right, and these will output two suggested wing root max bending moments to consider for input into the program.
After you have put in the max bending moment, you then start playing with the layup schedule to match the wing bending loads vs span, which is graphed (the top graph). Typically, if there is carbon fiber in the wing, it will see most if not ally of the load before the fiberglass sees any, so if you are doing a carbon wing, compare only the carbon curve to the load curve. What you are looking for, is for the cabon load curve to be completely above the aero load curve. The only exception to this is near the root, if you are using a two piece wing with joiner and tube. In that case, the carbon load curve be less than the aero load curve for the portion of span that has joiner/tube in it.
A few hints:
This program will quickly show the wisdom of having a layup biased towards having more material on the top surface compared to the bottom. Also, it shows the wisdom of using lighter material, so that you can do the material dropoffs with the strength following the aero loads more efficiently. The best solution is infinitely light
cloth with continuous dropoffs.
If the curve shows that you are "just under the aero load curve, you can make a note to add just an inch or two to the appropriate dropoff layer to compensate.
Note that the output span, wing area, and wing volume are for a single half of the wing.
This spreadsheet (or at least the underlying equations) has been quite accurate over the last 8 or 9 years for me in doing wing structural design, with a typical error of about 5% over. That is, it predicts slightly pessimistic with the estimated weights coming out a bit over the actual weights. I've not changed the equations to match the observed data better because I tend to like a bit of reserve weight in building new planes, as I always find some unknown that adds to my initial weight estimate!
Any comments, suggestions ? Des commentaires, suggestions ?
Send a message to Joe Wurts.
Copyright © 1998 R/C Soaring Magazine - Joe Wurts
No commercial use or publication (e.g. on other www or ftp sites, print media)
without a written consent.
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