Load Test Question

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proppastie

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early on I had not the learning experiences I now have......I looked at the original plans last night and and Irv or Bill have 2 false ribs where I only have air....their leading edge nose rib spacing is 7" with the false ribs where my spacing is 21".......
 

tdfsks

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How deep is the spar ? Do you have a sketch of the cross section of the spar at the root showing the web and flange configuration and thickness ?

One thing that concerns me in this design is that the spar flange does not have any stiffening of its free edge. If you look at say the spar in a Lazair or a Le Pelican ultralight, which both have spars made from sheet, the free edge of the cap has a lip (3/8" at a guess) bent on it to provide stiffening of the outstanding flange. You might be able to bend up an angle and slide it in under the existing flange to double the thickness of the flange and provide this stiffening. Cessna, on the other hand, use a bar riveted to the compression spar flange with a bulb on the edge (it is a custom extrusion) to provide the necessary stiffening of the flange against buckling.

However I would suggest we figure out what the root bending stresses are in the flanges before going any further with ideas for strengthening and assess this stress against the likely crippling stress of the flanges.
 

tdfsks

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Is the wing planform (span, chord etc) the same as the Carbon Dragon ? If not what is the span, root chord and tip chord.
What is the max takeoff weight and wing weight ?
Did you calculate a root bending moment ? Is so, what did you get ? What assumptions did you make about spanwise lift distribution (constant loading (proportional to chord) or something like Schrenk) ?
 

BBerson

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Looks almost symmetrical. I don't see much problem that it's upside down. Just keep loading until the 1" bar caps starts to column buckle. Release the load (lift tip) at each increment to see if those skin buckles return to normal. Then move to the next increment.
The only way to get minimum weight is with occasional failures.
 

proppastie

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Is the wing planform (span, chord etc) the same as the Carbon Dragon ? If not what is the span, root chord and tip chord.
What is the max takeoff weight and wing weight ?
Did you calculate a root bending moment ? Is so, what did you get ? What assumptions did you make about spanwise lift distribution (constant loading (proportional to chord) or something like Schrenk) ?
Wing and planform is the same as the Carbon Dragon
Hopping for 400 lb max gross, wing weight now looks to be say 140 lb with stiffeners. (currently 69 lb LH wing)
Loading was by area planform method Loading distribution by area of wing.
attached: EAA spread sheet for WS 8.25 with moments.....Loading spread sheet for sand bags. Moment of test is less because the wing gained weight.
 

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tdfsks

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OK ... let me chew on this as time permits.

First thing was to check the spanwise lift distribution to test the validity of the assumption of uniform pressure on the wing. I used the Schrenk method and the calculated Cl distribution is shown below (yellow line). I don't know anything about the airfoils or what the twist is in the wing. I assumed 3 deg washout at the tip. The airfoils don't matter too much provided that the root and tip airfoils have a similar angle of zero lift. Anyway due to the approx 3:1 taper ratio of the wing, the Cl is higher outboard then one would get for a rectangular wing (i.e. there is little in the way of an elliptical distribution) and so the Cl distribution is more or less constant. So the assumption of uniform pressure is not a bad one. A pity because if the spanwise distribution was more elliptical, the root bending moment would have been reduced. Not to be.

The next thing was to calculate the root bending moment. I used a simple spreadsheet I posted some months back in another thread on composite spar design. You can find the spreadsheet and theory in that other thread. The bending moment on the wing centre line is M = 77334 in.lb and the shear is V = 693 lb. If your wing attaches to the fuselage in the same way as the Carbon Dragon, we can use the moment and shear one station out from the root for designing the root of the wing and these are M = 68519 in.lb and V = 643 lb. These values seem to be a little bit lower than the moments in your spar calculation spreadsheet (M = 87686 in.lb and V = 813 lb) and I am not sure of the reason for this but I don't really understand where these numbers came from. My calculated shear force matches that in your static load test spreadsheet and so I think all is well.

My calculated test pressure load is 0.064 lb/in2 which is similar to the value in your static load test spreadsheet.

Metal Dragon M and V.pngMetal Dragon Spanwise Cl Distribution.png


The next step will be to calculate the section properties of your spar, calculate the stresses and then make some assessments about crippling stress, inter-rivet buckling and column buckling of the spar cap. A job for tomorrow.

What is the rivet spacing between the LE skin and the flanges on the spar channel and also the spar web and the spar cap ? What size rivets ?
 

proppastie

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What is the rivet spacing between the LE skin and the flanges on the spar channel and also the spar web and the spar cap ? What size rivets ?
Moment calculations was at a wing weight of 80 lb ......wing test calculations was for a wing weight of 138 lb (subtract the weight of the wing for loading?)

3/32 AD rivets, SS blind rivets in LE to nose ribs. ....Spacing in spar caps 3/4" Spacing in LE and skin to flanges 1"

Moment I of wing and LE wing station 8.25 post #23

1593699160436.png
 
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wsimpso1

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early on I had not the learning experiences I now have......I looked at the original plans last night and and Irv or Bill have 2 false ribs where I only have air....their leading edge nose rib spacing is 7" with the false ribs where my spacing is 21".......
Hmm, critical loads or stresses for buckling and panel stresses go with the inverse of span squared. A span 3 times longer means critical load is 1/9 as big... more reason to believe that you need those false ribs or some other sort of stiffener.

Billski
 
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tdfsks

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Moment calculations was at a wing weight of 80 lb ......wing test calculations was for a wing weight of 138 lb
I used a wing weight of 70 lb / side or 140 lb total. So 400 - 140 = 260 lb lift load causing bending which you can see in my calcs.

I will post some further calculations of stresses and critical buckling stress later .... half done .... been a bit busy today. Stresses do not look unreasonably high though. The buckling is the challenge.

Looking at your photo's, I can see that the LE torque box does not react its load directly to the fuselage (i.e. no fitting at the leading edge). The Carbon Dragon is the same. It has a diagonal drag brace behind the spar which would help to pick up the torque load and react it as a vertical shear load at the rear wing attach. Also I guess some of this load will be reacted at the spar attachment by forward and aft loads in the top and bottom attachment fittings. How is the torque in your D box reacted ? Do you have a drag brace similar to the Carbon Dragon ? I need to understand this before commenting on the buckles in the leading edge skin at the root rib.
 

tdfsks

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I have calculated the stresses in the root of the wing. Both limit (5.33g) and ultimate (8g) stresses are shown. I calculated the best case (i.e. full section, except leading edge skin, effective) and then the worst case (i.e. all sheet metal buckled at low stress with the caps carrying all the load). The stresses don't look to be too bad for instance, if the material was 6061-T6 (weakest structural alloy), Fty = 35,000 psi and Ftu = 42,000 psi (from MIL-HDBK-5F). So even the ultimate stresses would be just below yield. See calculations below.

Next step is to calculate the buckling. For that I need to know what the materials are. What alloys have been used for the caps, web, angle and LE skin ?

IMG_8663.JPGIMG_8664.JPG
 

proppastie

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Also I guess some of this load will be reacted at the spar attachment by forward and aft loads in the top and bottom attachment fittings. How is the torque in your D box reacted ? Do you have a drag brace similar to the Carbon Dragon ? I need to understand this before commenting on the buckles in the leading edge skin at the root rib.
More pictures build log. White parts are steel, Drag Spar steel 4130 1/2 x .035 wall (span 20" to center support, tube chart says 750 Lb/per tube for that length) same size with rear spar fitting, and .062 flat steel..... attached to rear spar with 1/8 6061-t6......Caps on rear spar are 7075-T6.....Outboard attach drag spar .062 7075-T6

IMG_1166.JPGIMG_1169.JPGIMG_1167.JPGIMG_1171.JPG
 
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wsimpso1

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Hmm, tdfsks is doing a lot of professional engineering services for free here. I would say you owe him at least a bottle of good scotch over all this ...

I reviewed his calcs and they look correct. They do make the assumption that when the skin gets to buckling, we lose the contribution from all of the thin sheet metal. I would expect that you buckle either skin or angle on the compression side first, shifting the neutral axis away from the compression side and increasing y in the My/I calc, which further increases compression stress on the remaining elements. Doing these cases individually and calculating the critical stress for crippling in each might show you which element buckled first, and then the progression of the buckling.

If it turns out that the skin buckling is first, the false ribs might really be needed. They also might be needed if the skin buckling deforms the angle and drives its buckling early. If the angles buckle first or closely behind the skin buckling, they will need their critical load increased. Knotty little problem you have here.

I am increasingly leaning toward needing both skin stiffening and stiffening the angles. Skins are easy, those false ribs will go a long ways at low weight. For the angles, you can postulate several schemes and analyze them for crippling stresses. Thickening the angles and adding a return on them (making the angles into channels) look like good bets for improvement on crippling calcs. How to add them to the test article now that it is built? Hmmm, you could bend another set of angles and rivet them to the flange part of the angles along with the skins. This will add more at the flange, so look close at the calcs when deciding exactly how to execute this added starter.

One other thought here, if precluding crippling gets complicated and/or heavy, there is another path. Making those spar caps beefier will reduce compression strains and thus reduce compression stresses in all of the elements that are buckling. A search of the design space may show that you can preclude buckling within your flight envelope at low added weight by a combination of these changes. Remember also that beefing you do near the root may not be needed as much as you go out the span, so you might not need to carry these weight increases very far outboard.

Billski
 

plncraze

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Thanks for sharing this stuff guys! For those of us who fantasize about doing this ourselves you are providing some great examples. I have a few books on loads and stress and there are no "plug and chug" stress analysis worksheets out there LOL! When you share your experiences it helps all the rest of us a lot!
 

tdfsks

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The problem I have with including any of the LE skin in the calculations, other than the narrow strip that is riveted to the flange of the spar, is that when the bending load in the skin gets to the root rib it has nowhere to go. So any bending load in the skin, in the area of the wing root, must be transferred from the skin into the spar and this implies that the bending stress in the LE skin at the root rib drops to zero. If this is the case, there is a fairly complex state of stress in that skin in the root area. Billski, I agree with your suggestion on how to treat the skin in the analysis if we are analyzing a section a few feet outboard of the root rib (and assuming there were enough leading edge ribs and/or spanwise stiffeners to maintain the shape of the D box skin on the tension side), but in the bay between the root rib and the first rib outboard, where the buckling is occurring, I think it safest to ignore any bending contribution of the D box skin.

I have a suspicion that the buckle in the skin at the root might be due to some sort of warping of the D box because of the lack of restraint of the skin by the root rib but I need to think that through .... one step at a time. I also have some concerns about the strength of the root rib in general including the main section behind the spar but we will also get to that in time. Before that, we need to figure out the load path for the torsion load in the D box and how this is reacted into the fuselage.

I think the next step is to sort out what to do about the flanges of the spar that are buckling and to make sure that there is enough column strength in that spar, otherwise more nose ribs or stiffer web stiffeners might be required to stabilize it.
 

proppastie

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Hmm, tdfsks is doing a lot of professional engineering services for free here. I would say you owe him at least a bottle of good scotch over all this .
Absolutely except perhaps he does not drink, in which case we need to think of something else.
 

proppastie

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I think the next step is to sort out what to do about the flanges of the spar that are buckling and to make sure that there is enough column strength in that spar, otherwise more nose ribs or stiffer web stiffeners might be required to stabilize it.
Locating a stiffener forward the spar as close to the spar is the only practical solution ....with other stiffeners spaced out forward......post 29 is first cut at a design solution.....

Spread sheet post 33 shows stiffener crippling calculations and 2"calculations and 18" unstiffened span plate (for skin of LE) Bruhn C7 has more information.

FEA of assembly (less ribs) at 40% load and Limit load with stiffeners.


40%loading.pnglimit load with stiffeners.png
 
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BBerson

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There might be more web shear deflection than expected because of the tractor mount up? In flight the wing shear transfers directly into the opposite wing.
Mark Calder had a D-cell test failure in a similar set up on his blog.
 
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