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RE: Wing spar calculations 12 Dec 2013 00:15 #702

Jim,

Many thanks for taking the time to help me over the past week or so. I'm pretty happy with my understanding of the calculations involved now and hopefully this is the last time I'll bother you before placing an order for carbon rod.

You mention (below) to use a max cap stress of 200,000psi to prevent compressive buckling - but on page D-11 you are using 150,000psi (and 280,000psi on pg. D-10) to decrease the tip deflection. What constitutes an acceptable tip deflection? On page F-11 you suggest 0.25" per foot of span per g (presumably ok for the Monarch, under discussion). I assume I'll be safe if I calculate the tip deflection for the original spruce/plywood spar and then re-design for carbon rod caps and carbon fiber shear web to give an equal or lesser tip deflection?

Can you give me an idea of the cost of .092" x .220" rod per yard or per meter?

Thanks again,

Philip

PS - I don't suppose you have the shear / compression / tension strength figures for 1/32" ply handy (or a link to a tech. spec.)?



From: Jim Marske [This email address is being protected from spambots. You need JavaScript enabled to view it.]
Sent: 06 August 2011 19:56
To: 'Philip Lardner'
Subject: RE: Wing spar calculations




This isn’t your wing but you can get the idea of how the shear stress is derived.

I don’t have much free time to analyze your efforts but I see you are on the right track. Too bad my work is not in metric, right?

Oh, for max cap stress use 200,000 psi due to compressive buckling.



I have seen the Carbon Dragon fly and it was very impressive to watch. Jim Maupin was a little guy and weighed like 68 kg. He wanted me to fly his bird but the glider sure looked fragile. Steve Arndt built an all carbon version of it for his weight using the .092 x .220” rod. It also looked great in the air. Don’t remember how many rods he used in his spar.



From: Philip Lardner [This email address is being protected from spambots. You need JavaScript enabled to view it.]
Sent: Thursday, August 04, 2011 7:39 AM
To: 'Jim Marske'
Subject: RE: Wing spar calculations



Hi Jim,



In running the numbers for the Carbon Dragon wing through your formulae I've created the attached spreadsheet to try and calculate the loads and stresses on the wing and come up with a design using carbon rods and a carbon shear web to replace the original all-wood design (with CF tows in the caps). I'm reasonably happy with what I've done so far (though I haven't double checked everything yet). However, I've come a bit unstuck when trying to calculate the weights and loads on the shear web...



On page D-7 of your manual you state

"Vertical shear is simply the weight on the wing outwards of the point of interest... To find the shear web Vertical Sheer anywhere along the wing find the load on the wing outboard of the desired location."



Vs = W1 / h1.

What I can't figure out is how to calculate the weight (W1) of the wing outboard of any given point (grey panels at the bottom of the spreadsheet.)



All the best,



Philip.





From: Jim Marske [This email address is being protected from spambots. You need JavaScript enabled to view it.]
Sent: 01 August 2011 01:11
To: 'Philip Lardner'
Subject: RE: Wing spar calculations

You got me on that one. You are correct. Looks like I have some corrections to make. I’m surprised no one noticed it before. Congratulations, sharp eye!

I’ll make corrections on D-5 tapered wing also.





From: Philip Lardner [This email address is being protected from spambots. You need JavaScript enabled to view it.]
Sent: Sunday, July 31, 2011 2:58 PM
To: 'Jim Marske'
Subject: RE: Wing spar calculations



Thanks Jim - I knew it was something obvious!



Next problem - on page D4, Finding the moments along a semi span of a tapered wing, you show the centroids for the four panels as 96.9", 74.7", 51.8" and 27.4". However, when I use the formula you give on page D1 for finding the centroid of a single-tapered wing (L * (2 * Ct + Cr)) / 3 * (Ct + Cr) I am coming up with slightly different numbers for each panel: 100", 77.14", 53.33" and 28". What am I doing wrong... or am I using the wrong formula?



Thanks again,



Philip.





From: Jim Marske [This email address is being protected from spambots. You need JavaScript enabled to view it.]
Sent: 30 July 2011 20:02
To: 'Philip Lardner'
Subject: RE: Wing spar calculations

The difference between pages D3 and D4 is one wing is rectangular and the other is tapered.



The area of the 40 x 60” element on page D3 is in square inches, but we have to convert to square feet.

There are 144 square inches in a square foot. So, 60 x 40 = 2400 square inches / 144 = 16.67 square feet.



On page D4, find the area of the tapered wing.

Root chord + tip chord / 2 = average chord. So, 60” + 20” / 2 = 40 inches average chord.

Now, 40” x 240” = 9,600 sq in / 144 = 66.67 sq ft.



Jim M





From: Philip Lardner [This email address is being protected from spambots. You need JavaScript enabled to view it.]
Sent: Saturday, July 30, 2011 2:39 PM
To: This email address is being protected from spambots. You need JavaScript enabled to view it.
Subject: Wing spar calculations



Hi Jim,



I am working my way through your Composite Design Manual and have run into a problem.



On page D1 you describe how to find the centroid and bending moment for a wing treated as a single panel:



Centroid = (L * (2 * Ct + Cr)) / 3 * (Ct + Cr)



Moment = Gross Weight / 2 * Centroid x 8g



...all very simple to follow.



However, on page D4 (and D3), "Finding the moments along the semi-span of a tapered wing planform" you appear to change the formula to:



Area = (((Cr + Ct) / 2) * L) / 144 ==> (((60 + 20) / 2) * 240) / 144



I can not see where you are getting the '144' from or what it represents. I'm sorry if I'm being thick (or blind!) but I need to understand the working if I'm to apply it to my own project.



No doubt I will have a few more questions down the line, but for now any help with this problem would be much appreciated.



All the best,



Phil Lardner

Ireland.
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