Wing Bolt Shear Strength

Knowledge based discussion here... Bored at work and started thinking about those 16 bolts that my life depends on. On the -7 we have 8 NAS1307 & 8 NAS1304 bolts rated at a min shear strength of 95,000 psi & a tension strength of 180,000 psi... Has anyone been nerdy enough to figure out what types forces are placed on this area at 6 g's? While the wing attachment area is small, it's quite beefy when compared to big c-310s & c-406's. (Planes much larger & heavier)... I'm definately not questioning the strength of our wing attachement areas. I'm just kind of curious what types of forces are placed on these bolts... The pressure must be IMMENCE for this 4-6inch area!!! So, does anyone know what actual structural strength of the wing root is vs. the flight envelope? Obviously, I'm not an engineer Just random thoughts.

Adequate

No I haven't calculated the ultimate safety factor of the wing attach bolts (although I could). They are adequate for the application. How do I know? Because Vans load tested an actual wing to the 6 g's plus the 1.5 safety factor above that. The wing and wing bolts didn't break. One test is worth a thousand opinions.

Numbers please.

NUMBERS! We don't need no stinking Numbers.

I could calculate it, but if I start doing that i'm going to end up getting dragged into designing the tapered wing that i've always wanted for the -6. I don't have time for that yet, so i'm going to resist temptation... For now.

Taperwing on a 6 from snowflake

You just have to tease us, I really think you should proceed on your dream,
besides I would hope you would share with the rest of us so you are not the only 6 out there with tapered wings --------------we have to have a little fun
Ken
Rv6 flying
assembling a rocket

go look at the spar/supports on say a decathlon and then tell me the RV looks weak

seeing finishing nails and wood spars on an aerobatic plane, scares me.

seeing finishing nails and wood spars on an aerobatic plane, scares me.

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You only have to worry when the nails start backing out. If you replace them with pole barn spikes you never have to worry again.

Just doing the napkin math, and working on dimensions from memory.

Assumptions:
1800 lb gross weight split equally between each wing
6g load
The load on the wings acts 60" out from the joint.

We're now looking at 324,000 in*lb bending moment at the joint.
Assuming the spar is 9" deep:
Each web is supporting ~72,000 lbs of shear force.

Edit: forgot to distribute between the 2 webs as pointed out below, 36,000 lbs is the correct value for shear force.

How much each of the 4 bolts (in each web) is taking is more than I can tackle off the top of my head. Here are the conditions I see though. They are loaded in double shear and they take a varying amount of load decreasing towards the center. The smaller bolts will deform slightly more, and pass their load off to the main bolts. The designers probably did that to allow stress to flow around the small bolts to the larger bolts efficiently. Ideally the joint is designed such to allow enough elastic (non-permanent) deformation to load all fasteners equally relative to its strength. Often the limiting factor in a joint like this is actually the bearing strength of the aluminum around the joint. It's got to absorb the load of that strong hard steel bolt.

If I get a chance, I'll check my assumptions and dig into the details later.

Guy

Good Stuff

Guy... Now, that was what I was looking for! Forgive me if this is a dumb question. But how do inch/lbs translate to PSI? The Bolts have a minimum shear stength of 95,000 PSI... I'm guessing the ultimate outcome is that the spar will fail else where WAY before anywhere near that joint!!! I'd like to get my hands on the info from the static tests van's did. All i've found are different bits like The spar can withstand +9 Gs for 3 seconds before failing... Thats 14,400 pouds of force on that sucker Geeeezzz!!!!

bkthomps said:

seeing finishing nails and wood spars on an aerobatic plane, scares me.

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Don't ever set foot in a Pitts then.

Guy, if your distances are correct, then your numbers for spanwise spar cap load are 2X high, so I'd call it: "36,000lb tension in the spar cap", which is reacted as shear in the 4 bolts in each cap. (do a moment balance at the root of the lower spar and you see that the load in the upper spar is 36kip compression, which means it's 36kip compression in the lower, basically, you divided the moment by 4.5", not 9")

Usually you can assume equal STRESS (not load) in all 4 bolts as a starting point in joints like this, so I'll add the area of the 4 bolts: Pi*(.125^2+.219^2)*2 (of each diameter) *2 (double shear) = .79 in^2

36kip/.79in^2=45.5ksi in the bolts, or a 2:1 margin at 6g.

Guy is correct that the aluminum either in tension, compression, or bearing is likely the limiting factor of the joint, but I'd have to have a wing spar or a drawing of it around around to do that.

For fun, IF the spar cap is a piece of 1"x1" bar there, then each Of the 7/16" bolts would have ~6900lb on it, acting on a bearing area of 1" * .4375" = .4375 for a bearing stress of 15,000psi. 2024-T4 (a guess..) has a bearing yield allowable of 64000psi, so over 4:1 factor of safety for causing permanent damage, 5 or 6:1 for actual failure (assuming a 2 D edge distance).

Again, IF the cap is 1x1, then the tensile stress in the cap is 36000lb /1in^2 or 36000 psi, with a yield allowable of 47000 for a 1.3 factor against permamant deformation.

The other issue is that the two 1/4" AN4's in the spar web that were the subject of the SB will in THEORY carry almost all the vertical shear load (assuming they're installed). For that you do:

900lb*6g=5400lb

Area of both bolts: 2 (double shear) * 2 (bolts) * Pi * .125^2 = .2 in^2

Shear stress: 5400/0.2^2 = 27000 psi. I think the shear for AN bolts is 75ksi (75000 psi), so you've got ~3.5 margin, and that's conservative because the other 8 bolts take SOME of the vertical shear.

Of course all that's conservative because the weight of the fuel and wings is outboard of that joint and actually subtracts from most of those stresses.

Guy Prevost said:

Just doing the napkin math, and working on dimensions from memory.

Assumptions:
1800 lb gross weight split equally between each wing
6g load
The load on the wings acts 60" out from the joint.

We're now looking at 324,000 in*lb bending moment at the joint.
Assuming the spar is 9" deep:
Each web is supporting ~72,000 lbs of shear force.

How much each of the 4 bolts (in each web) is taking is more than I can tackle off the top of my head. Here are the conditions I see though. They are loaded in double shear and they take a varying amount of load decreasing towards the center. The smaller bolts will deform slightly more, and pass their load off to the main bolts. The designers probably did that to allow stress to flow around the small bolts to the larger bolts efficiently. Ideally the joint is designed such to allow enough elastic (non-permanent) deformation to load all fasteners equally relative to its strength. Often the limiting factor in a joint like this is actually the bearing strength of the aluminum around the joint. It's got to absorb the load of that strong hard steel bolt.

If I get a chance, I'll check my assumptions and dig into the details later.

Guy

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sheepdog said:

besides I would hope you would share with the rest of us so you are not the only 6 out there with tapered wings

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Sadly, I wouldn't even be the first. I'm aware of at least two, maybe three, -6's built with tapered wings. I think one has retractable gear, too.

Reverse-engineering spar joints

I have reverse-engineered the spar joint on the RV-7/8 very extensively, and also the RV-6. This work was done for a couple of interesting projects that are ongoing.

Some of Guy's dimensions are way off, but the thought process is about right. The spar cap centerlines are 6.25" apart, not 9". The center of lift is at about 45% semispan for the rectangular wing, but you have to make some good assumptions about how much lift is carried over onto the fuselage.

gtmule's numbers are pretty close to mine, there is about 36,000 lbs of tension and compression in the spar caps at the point where the bolts start picking up load. This is based on 1600 lb aerobatic weight at 6 g's. Also included is the added lift to offset the tail download to trim at 6 g's at Vc (the highest speed for which a 6 g maneuver is allowed)

The numbers come out fairly different depending on how much lift you assume is carried over onto the fuselage, and what you include for trimming loads, and how/where you assume the fuselage weight is "hung" on the spar. These uncertainties are part of why we maintain safety factors of 1.5 in design. The bolts themselves do have a SF = about 2 (and that is appropriate, there is language in part 23 about bolted joints having extra SF over and above the 1.5)

There are special methods for determining the load sharing between several bolts in an assembly, and it is often a poor assumption that they all share equally. Using the smaller bolts in the first and last holes does help to even the load sharing out, and that's why they do it.

I did check the bearing stress in the aluminum caps and carry through bars and it is not the limiting issue.

What is the critical issue is the tensile and compressive stress in the spar caps and the carry through bars alongside the bolt holes. The stress here depends on what you assume for stress concentration factors, and how many cycles to that stress level you want the structure to survive without fatigue. Making some reasonable assumptions here shows that we have adequate service life.

I can tell you that the critical components are the carry through bars, they are somewhat smaller in net cross section than the spar caps. You can assume that the bulkhead webs will help share some of the load, but that is hard to substantiate without test. (And yes, tests have been done by Van's, and we are OK)

Using Mil-hdbk-5 data for material properties, I can tell you that the spar joint is just about exactly the right strength, and not hardly any extra. I have found that mil-hdbk-5 is usually conservative, and material suppliers (like Alcoa) will quote somewhat higher properties, which gives us a bit more margin. That's nice. But I would never advise exceeding Van's recommended weights and g-loads after doing this study.

A bit about the RV-6. There are a few folks on the forum that have stated they believe that the RV-6 is very strong, even over-built in the spar area. They base this on their instinct looking at all the bolts holding it together. But the RV-6 spar caps have smaller net cross sectional area than the -7 and -8. And the bolts that hold the center splice plates to the ends of the spars are smaller. In particular, the 1/4" bolt in the splice plate has a lower safety factor than the other bolts.

The result: again, I think the spar assembly is adequate, but very little extra. I would not ever accept an increase in aerobatic weight or g's over what Van's has recommended.

I believe there are a number of RV-6's that have had builder-selected gross weights above the Van's factory values. I would advise caution on those airplanes - keep your g's under control.

scsmith said:

There are special methods for determining the load sharing between several bolts in an assembly, and it is often a poor assumption that they all share equally. Using the smaller bolts in the first and last holes does help to even the load sharing out, and that's why they do it.

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Agreed, If it were 4 equally sized bolts, I would probably suggest ignoring the middle two for a conservative "back of the envelope" analysis, but since they're so much difference in the diameters/areas/stiffnesses between 1/4" and 7/16" (3:1 area ratio) I think it's a reasonable bet that all four of the bolts are under something like the same STRESS (not load).

Its really the kind of thing that's best measured in the lab on a mockup of the joint, because there's 10 different hand calculation approaches, and 10 different finite element approaches that all make different assumptions about the effects of friction, preload, out of plane stiffness, etc. Do you have actual dimensions for the spar caps and carry-through caps? I might run the test for fun one day after work....

It's my experience (in a test lab, and designing parts) that the mil-hdbk-5 (now MMPDS) static properties data for the older alloys (7075, 7079, 2024, 6061) are very conservative, because of all the improvements in material processing since a lot of that data was developed. Some of the newer alloys have more recent data, so that modern processing is already factored in. Of course with 7075, 2024, etc, the designer can't take credit for that extra margin with the FAA or Military Branch, but he knows it's there, and Poisson, Newton, Hooke, etc are all secretly giving him credit....

edit: Of course MMPDS/Mil-hdbk-5 data is also all based on confidence intervals (A-basis, B-basis, etc,) so it's really intentionally conservative, where as manufacturers can report the raw average of their test sample....