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large hole through main spar center section

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raymondbird

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Need to get air through my center section for rear mounted radiator. As I understand the attached drawing represents the shear load across a one piece wing spar, (zero across the fuselage) so does that mean a large full height hole between the spar caps would be ok?

Any knowledge or opinions greatly appreciated! Shear.gif
 

mcrae0104

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The diagram generically represents the shear loading if the shear loads are transferred to the fuselage at two points. Assuming that is the case with your plane, then yes, a hole may be possible as long as the moment can be carried all the way through the center of the spar across the opening you need. More details such as the type of aircraft of a drawing of the center spar section would help. Is it wood, metal, or composite?
 

raymondbird

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Your diagram probably represents the fuselage side walls pushing on the spar.
The drawing came with this:
The diagram generically represents the shear loading if the shear loads are transferred to the fuselage at two points. Assuming that is the case with your plane, then yes, a hole may be possible as long as the moment can be carried all the way through the center of the spar across the opening you need. More details such as the type of aircraft of a drawing of the center spar section would help. Is it wood, metal, or composite?

Great! It is a wood core spar with pultruded carbon rod caps (.22 x.092) 24 of them top and bottom and many layers of biaxial cloth shear web.

Oh, and here is the text that came with the above shear load drawing. It's from a man named Kenney Anderson:

On the top is the distributed load on the wing. Nothing is to scale. The middle diagram is the shear diagram. The shear is the maximum at the root, and is equal to the total vertical load coming from the wing. Because there is an equal load being reacted at the fuselage, those 2 loads balance (wing load up = reaction down at the fuselage). Therefore, we see that there is no new shear being introduced between these two attach points so the shear in the middle is zero (implication is that you don’t need much of a shear web (so you can tie just the upper and lower caps together).
SDC10798.JPG
 

raymondbird

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The diagram generically represents the shear loading if the shear loads are transferred to the fuselage at two points. Assuming that is the case with your plane, then yes, a hole may be possible as long as the moment can be carried all the way through the center of the spar across the opening you need. More details such as the type of aircraft of a drawing of the center spar section would help. Is it wood, metal, or composite?
Oh yes, it is bolted to the fuse at two points.
Thanks
 

mcrae0104

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With no shear between the two attachment points, you could analyze the caps as two-force axial members (columns) and disregard the shear. But as BoKu pointed out, the shear will not be zero when the load on each wing is different. I suggest you construct shear and moment diagrams for these cases (positive and negative limit load, plus reasonable assymmetrical loads) so that you can determine how much of the web can be removed between the attach points.
 

mcrae0104

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The web still supports the caps from vertical compression. Might want some vertical stiffeners. I would keep the hole minimum size with flange added.
Please define what you mean by vertical compression.
 

BBerson

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Please define what you mean by vertical compression.
Could be called vertical shear, I guess. Vertical shear is rarely mentioned. While horizontal shear cancels across to zero at the fuselage, I think vertical shear is maximum. The upper cap is in column compression in flight and without a sufficient web between the attach points the upper cap would collapse down (buckle).
A 1" hole would hardly matter. But the OP asked if a "full height" hole would be ok. Presumably large enough for a radiator duct.
 
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wsimpso1

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More complicated question than you think. If symmetric pull up were the only load, yeah, you would just have to convert bending moment at the root into tension in one element and compression in the other, then make the elements elastically stable to these loads - yeah, they are not just caps, they become beams themselves in the designs I know of.

Then there are asymmetric loads. One wing making different lift from the other due to turbulence, control surface deflections, etc. I would have noodle on that a bit as asymmetric loads mean different shear and moments input to the mounts from each side... Draw a free body diagram for the carry-through as two beams and start your own noodling. But those new members between the mounts will have to carry loads independently...

In many cases, the designers just bend the entire beam around the duct in question. It sure makes changing jet engines easier that way. I promise I will noodle on this some more, but I promise no remarkable gain in wisdom on this topic.

Billski
 
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flyboy2160

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Even though the vertical shear loads are taken by the fuselage (assuming it is much stiffer than the wing spar), the center beam section still must carry the high root bending loads across between the fuselage. The center shear webs should be analyzed to make sure they won't buckle from this bending load. If you do the calcs, you'll probably find that, by far, the largest part of the principal stresses are these bending loads.
 

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wsimpso1

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OK, worked on the wing tip molds and the cabin air mixer box, and got to a stopping point, then came back in a looked at this again.

Assuming the wing and fuselage are solidly mounted to each other at the fuselage walls, the wing shear is dropped into the fuselage at the walls, while bending moment passes through the fuselage to the other wing. In a symmetric pull with equal shear and bending from each wing, the problem is simple - the carry-through sees only bending. The bending can resolve to simple tensile and compressive loads if the web is omitted from the spar, but additional bending is present. The curvature of the spar just outside of mount can be expected to carry through the center section as well. This curvature = M/(EI) just outboard and will result in bending stresses in each cap, and must be superimposed upon the tensile/compressive stresses. Tension is fairly straightforward, but compression will require that buckling be precluded using curved column theory. I do not know about the Aero structural texts, but us Mechanical Engineers use Roark's for this sort of thing.

Go through FAR Part 23, and you will see that we also have to cover a certain amount of asymmetric lift. The shear through the fuselage will be the difference in lift between the two wings. The moment imposed upon the root section will not be constant, but will increase from the low lift wing toward the high lift wing. The moment at the high lift wing will be greater than the low lift by the amount of the shear in through the fuselage times the width between fuselage walls. Then we carry wing curvature from just outside the fuselage through the fuselage.

So, for the spar through the fuselage, starting with asymmetric lift, you need to calculate the shear and bending moment diagrams, Through the fuselage, the shear is the difference between the lift of the two wings, and highest moment is highest moment in the high lift wing. The spar devolves into two elements, one with tension, one with compression. Load summary:
  • Magnitude of axial loads is Max moment divided by distance between cap centroids;
  • The shear difference between the wings must be carried by the set, but since shear sharing may be difficult to compute, I suggest that each element should be designed to carry all of the shear;
  • The wing curvature at the root will be carried through the axial elements through the fuselage, with the resultant stresses being superimposed upon the axial stresses.
  • Spar fuselage elements will also be subject to curved column failure and must be precluded.
Ends up being a tad more complicated than might have been hoped. Maybe a spar curved around the duct will be more compact...

Billski
 

Dana

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My Fisher 404 had a hole through the lower center section spar, not on the original design. The plans called for the elevator pushrod to run under the spar, exposed under the bottom of the fuselage. The builder moved it up inside and ran it through the spar, with some additional reinforcement. It was certainly a much cleaner installation though it made me a bit nervous, but the plane had accumulated around 300 hours by the time I bought it so I didn't worry too much about it.
 

blane.c

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Back when I worked in home-building some of the designs were for electric heat and when those designs traveled north (to colder climates) some builders elected to go with hydronic baseboard heating which was a good substitute. Others went with forced air and without dedicated duct routing in the plans when the sheet metal "squads" left the job site there was all manner of wood debris littered about the basement and "soft spots" in the floors above. Just saying I haven't had good experience with unplanned duct-work going through structural support areas.
 

raymondbird

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Back when I worked in home-building some of the designs were for electric heat and when those designs traveled north (to colder climates) some builders elected to go with hydronic baseboard heating which was a good substitute. Others went with forced air and without dedicated duct routing in the plans when the sheet metal "squads" left the job site there was all manner of wood debris littered about the basement and "soft spots" in the floors above. Just saying I haven't had good experience with unplanned duct-work going through structural support areas.

Okay thanks again guys! Not as easy as I thought eh. Obviously WW2 fighters had all manner of holes thru the spar for cannons and stuff, but of course, we don't have teams of engineers working for us. The Bell P39 even had giant cooling ducts but it used some sort of steel tube truss center section.
 

BBerson

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Perhaps a triangular opening (basically making a truss between the attach points) would be a workable solution too.
Yes. The entire center web could be removed and replaced with a single cross diagonal or X brace.
It might be easier to make the entire center truss with steel tube with wing attach fittings welded on.
I think the Beech 18 might be similar.
 

fly2kads

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Yes. The entire center web could be removed and replaced with a single cross diagonal or X brace.
It might be easier to make the entire center truss with steel tube with wing attach fittings welded on.
I think the Beech 18 might be similar.
That's the route that VanGrunsven took with his RV-1 (modified Stits Playboy). The new cantilever wings attach to a steel truss across the fuselage.
 
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