Spitfires…..Why Are Their Wings On Upside-Down?

Thank you Vega ECM for some ‘proper’ engineering thought, based upon experience. In my perfect forum, this stuff would just keep on coming.

Interference fits are not used to encourage frictional load transfer, quite the opposite in fact. An interference fit represent the best standard of mechanical i.e. metal pressing against metal, contact with the minimum practically achievable risk of slippage. Note that contrary to popular belief an interference fit is not a fully uniform contact in that it’s actually a series of contacting patches at the surface e.g. an interference fit bolt will leak fluids. So when loaded there will still be tiny movements, some of which will cause metal to metal slippage and this will damage the surface, but the interference fit has another advantage to deal with this. An interference fit joint inherently creates a region of compressive surface stress at the interface. Hence if areas surface damage causes cracks to form, then these will be held closed together by the compressive stress. A crack which is held closed will not grow.

Overall an interference fit gives a good mechanical load transfer, allows practically the minimal degree of deflection which promotes even load transfer in multiple fasteners/members, is robust to coping with its own damage and absolutely maximises fatigue life.

Thank you Vega ECM. It has certainly made things clearer for me.

I would ask you to clarify, particularly bearing in mind your comments on friction not to be relied upon, the Spitfire’s multi-tube spar? And perhaps the use of interference fit bolts for attaching the wing to the fuselage. I grant that the latter will not be aluminium meeting aluminium, and probably the former isn’t either, but presumably both are partially relying upon friction?

Has anyone considered the strength of the bracket, ie more land on the lower bracket above and below the smaller holes?

I reckon this thread needs to consider some basic beam structure theory and how this is applied in Aircraft Design;-

General Bending Beam Theory – A symmetrical beam in bending (as the Spitfire is) will experience an equal, but opposite stress in in the top and bottom booms (i.e. the bits what the holes are in)

Defined by Outer fibre Boom Stress = Bending moment x Distance from centre / section inertia

Hence it makes no difference if it’s up or download bending the boom stresses top & bottom are the same. So I’m afraid any suggestions that the + 6g one way & – 3.5g the other or which side of the joint has the Landing Gear attached has nothing to do with why it looks as it does.

Aeroplane Stressing Rule 1 – In a joint where you have to use multiple fasteners, your uncertain as to how much load each one is actually carrying (aka structural indeterminacy or Hyper-static’s)

So what I do is try to use the largest number of the smallest diameter within the allowable space while respecting minimum cross section rules. By doing this I’m trying to encourage any single fastener, which as a result of build tolerances could tend to become preferentially loaded, to elastically deflect and hence this movement will take up any gaps on the others, as such in encourages load transfer on to other the fasteners……I’m trying to promote an even load distribution on to all the bolts

So this is why there’s four smaller holes in the more spacious bottom boom and three larger holes in the top where the space is more restricted. Both will work but the lower boom is better design practice.

As for friction;-

Aeroplane Stressing Rule 2 – In the real world, Friction within structures is unpredictable;- so it can never be depended upon for structural safety but equally it will always show up when you don’t want it to.

So if I have friction, which might be beneficial I deliberately ignore it, making a safe structure purely with mechanical means, but if I know I could have friction which is detrimental I design the structure to be safe with a very conservative assumption (typically 0.8).

Also producing a design which clamps two aluminium surfaces tightly together specifically for frictional load transfer is a really bad idea because the surfaces will gaul/fret (it’s just too soft a material) and resulting damage will seed fatigue failures.

This contribution comes from my mis spent 25 years in aircraft design;-.

But the G forces acting upon the wing in combat and pulling positive G will be acting the opposite way, I.e trying to push the wing down …….

So what is supporting the other masses of the aircraft, e.g. fuselage, pilot, engine….? It’s usually referred to as lift.
Keith

Yes aluminium alloy CD. TonyTs lovely picture shows the small plate that states the size of the bolts that are fitted. These plates often have the aircraft serial number on and have been used as evidence of ID for a rebuild.

Well, not the wings exactly…..but why are the wing mounting bolts positioned upside-down?

This photograph by David Whitworth (which I hope he doesn’t mind me using) illustrates my point perfectly.

Basically, why are the larger bolts on the upper spar? You’d think the larger bolts should be on the lower spar (because the lift under the wings and the weight of the aircraft on the undercarriage will push the wings up). In fact, why are the upper and lower bolts different sizes at all? They should be the same size. (Unless there is some other device preventing the lower spar’s tendency to pull away from the fuselage?)

At first I thought it was because there was some obstruction to prevent the lower spar’s tendency to push towards the fuselage, except the lower spar isn’t going to push towards the fuselage; it is going to tend to pull away from the fuselage! (And the upper spar, which could be up against an obstruction, has larger bolts?)

So, if the upper and lower spar are held only by the friction between the spar surface and the mounting surface, which they are (apparently), and the friction is proportional to the tension provided by the bolts (which it is), why are the bolts different sizes at all?

Yes, there are three upper bolts and four lower bolts (again, why?) but the three upper bolts are much stronger than the four lower bolts. Why?

But the G forces acting upon the wing in combat and pulling positive G will be acting the opposite way, I.e trying to push the wing down and they do get reamed out on installation.

The wing spars are concentric square sections…

…and they are aluminium alloy?

As for the friction relief gained by the overtightening of the bolts I have not heard of that being employed in stress calculations…

For my money the joint is all about the friction; the bolts shouldn’t experience any shear force at all (unless they are not torqued-up properly)!

Actually, the more I look at the joint, the less I like it! No matter how tight the inner two (or three) bolts are, if the outer bolt is loose, the entire load on the wing is concentrated on the narrowest sections of the spar-boom above and below the bolt. If these fail all the inner two (or three) bolts will do is ensure the broken-off end of the spar-boom is found in the wreckage, still attached to the fuselage!

The carry thru spars are fitted to frame 5 which is adjacent to the firewall. The wing spars are concentric square sections which are diff to make but make sense to avoid crack propagation. Each wing has a lower and upper spar fitted around the qtr chord……bedtime as I am rambling!!

…the photo above of the lower spar boom with the 4 holes looks as if the web is thicker than the top boom.

I think that is a photograph of the through-fuselage ‘spar’ that the wing-spar booms bolt to?

The bolts do have an undercut. There are then washers of varying thickeness. Early spits had stainless bolts, later steel. The washers were stainless. I have several very corroded bolts from top and bottom kicking around. The washers are beautifully marked.

Question I cannot resolve from the internet is a) Do the bolts have a shoulder, is the washer back undercut? b) the photo above of the lower spar boom with the 4 holes looks as if the web is thicker than the top boom.
Now with my basic stress knowledge, I agree that the 3 bolts in the top boom are due to lack of linear space, and the strength of the top boom bolts could be less than the strength of the bottom boom as the stress would have been (I think) calculated on the basis of a 6g positive loading & a 3.5g negative loading times the military safety factor of 1.25 from AvP970 or equivalent in mid 1930’s.
As for the friction relief gained by the overtightening of the bolts I have not heard of that being employed in stress calculations. In any case as flight loads would be oscillatory in nature, & I think this action could negate any gain from friction, so relying purely on the bolt (double) shear strength & the spar boom holes bearing strength.

4 bolts would give a more gentle load distribution from one part to the other, so could also then be a smaller dia.

Another question then, with the ever increasing weight & performance of the Spitfire, was the spar joint overstrength in the first place or how was it strengthened
with each increase?
Interesting thread.
Keith

The shank (unthreaded) part of the bolts is the same length as the holes in the spars and the castle nuts have a fixed washer extending below the threaded area.

On the mk V the Tob Bolts are 0.9375″ and the bottom ones 0.71875″.

Thanks for providing the exact sizes for a comparison! 🙂

A cunning plan indeed (and basically what I was expecting to see had I not misinterpreted the original photograph); the combined cross-sectional area of the three upper bolts is 2.07 square inches, the four lower bolts 1.62 square inches, so reasonably close (and not taking account of the threaded portion). I guess we can approximate the cross-sectional area to the tension applied by the bolts and therefore the frictional forces at the spar / mounting interface.

There are no steel ferules in the wing attachments, but I think I’ve heard of a modification to allow their use as a “repair” for oversize holes; if you look at the shape of the fuselage at frame 5, it is wider at the top spar attachment point, and so less room for 4 bolts, no doubt a cunning plan was hatched which meant that the top three, carried the same load as the bottom four.

It may be that ferrules of steel are inserted into the larger holes but the bolt sizes are the same?

Not on the wing spars I think; the Spitfire spars (or should I say, spar booms) are the famous ‘solid’ aluminium box-in-box-in-box design.

Do you have photos of the actual spar pins or bolts used to make the connection. It may be that ferrules of steel are inserted into the larger holes but the bolt sizes are the same ? If I recall the engine bearer connection to aircraft frame for many Merlin applications used a steel ferrule to prevent crushing of the member being fitted.

Spar connections always amaze me. Knuckle joints for Navy aircraft with folding wings amaze me even more. Watching a Corsair fold down its wings then take off always gets a wince going on my face. All those forces concentrated into a folding joint !

Biplane spar connections are made using chicken bones, relying on the trussed box section bi-wing assembly to demonstrate its simple soundness.