Only the main spar went through the fuselage. ( see page 52) the wing skin below the fuselage was the tank bay cover.
cimmex
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Only the main spar went through the fuselage. ( see page 52) the wing skin below the fuselage was the tank bay cover.One obvious thing we've overlooked despite the photographs posted is the Fw 190s use of a carry through spars, there were no bolts and I imagine the lower wing skins were continuous across the fuselage.
Do try reading what I wrote; the only part of the Spitfire wing where wrinkles could be ignored, was between ribs 14 - 19, which is the where the aileron hinges attach to the rear spar. Any wrinkles, anywhere else, forward or aft of the mainspar, would necessitate bolts and wing inspection, with probable wing replacement.The concern for ensuring that overstressed Spitfire wings were wrinkle free on the leading edge but could tolerate with indifference 2.5mm wrinkles aft of the main spar is evidence again that the spitfire in concept is a single spar design.
Like four .303" Brownings, or cannon, perhaps?A two spar aircraft is almost completely indifferent to the state of the leading edge skins and the leading edges can be penetrated by landing lights, pitot tubes, gun ports without any serious engineering.
As said above, SEVEN bolts + ONE on the rear spar.I believe the Spitfire used no less than 4 bolts to attach the main spar. Evidence not only of the load this spar needed to carry but also of design to resist torsional loads.
Do a second thought experiment, and think what will happen to the wing, when the ailerons (attached to the rear spar) are operated, and the sliced-through rear spar is free to flap up and down, instead of moving the wing.Do a thought experiment; take an angle grinder to a spitfire and and cut from the trailing edge at the wing root to the main spar. The wings strength will be minimally effected even with loss of the rest secondary spar and continuity of the skin. Most of the lifting and torsional loads go through the front spar.
Nice to see that you agree with us.Do that to a two spar design and you've likely to loose the wing as you've lost the important rear spar and the thick upper skins that form a torsion box.
Main spas: 7 bolts per spar. 4 bolts per lower 'beam' (tube in tube construction; maybe people can remeber biology classes - the tube-in-tub 'construction' is applied by plants and bones); 3 bolts per upper 'beam'.
Here are photos:
http://www.ww2aircraft.net/forum/aviation/fw-190-roots-great-roll-rate-34677-3.html#post954157
For rear spars: 1 bolt.
The leading edge and main spar of a Spitfire Are violated with Cannon and Machine guns. Having said that penetrating a box of H or I section type spar in the center of the shear web is a fault that can be compensated for with suitable doublers to compensate for the loss of shear transfer from top to bottom cap in bending.
Nope the wing skins were no continuous from tip to tip. There was a large removable panel to give access to the fuel tanks. The rear spar was not 'carry through', it ended and was bolted to the fuselage.
http://www.albentley-drawings.com/images/FW190A6W.jpg
I accept that the rear spar bolt is critical in providing torsional stiffness; however I see the rear spar of almost total insignificance. It could have been eliminated.
If the rear spar bolt is there - and significant in size - then the cross section of the bolt will tell you how much shear stress it was designed to take - as well as some tensile stress. As to insignificant, It would have to match the bolt in load transfer.
Firstly there is almost zero torsional load on the spitfire wing when flying statically since the centre of lift is where the spar is at around 25% of chord.
Minor torsional loads would come from pitching moment as AOA varied; however aileron torsional loads would be more.
There are at least three primary contributors to Torsional loads from the wing which must be taken out in shear at the fuselage. First, the Pitching Moment of the airfoil due to the Aero loads on the wing, Second, deployed flaps, third - the asymetrical loads induced by the ailerons. Depending on whether the cannon recoil axis is on the Centroidal axis of the wing, it to will cause a torque about the Y axis in the wing. Each must be looked at for combinations of loads and stresses for both the load paths and the attach structure.
How much? 100kg per aileron maybe? That's only about 1000 Newton Meters. Not much really.
You don't have either the aero loads imposed on the rigid structure (first pass) or the revised combination loads taking into account XYZ deflections due to combined Axial and Normal Forces as well as introduced Moments discussed above. "Not much Really" is somewhat misinformed observation based on zero facts.
The torsional loads from the aileron would be transferred via the ribs at the hinge points to the box formed between the main spar and the leading edge which is part of that same rib. The ribs at the leading edge have elaborate triangular trusses so that leading edge box is much much stiffer than the main spar alone.
The Moment created about the Y axis of the Wing is the X axis force applied as flat plate pressure from the freestream flow upon the aileron. The aileron structure in turn transfers the load via the hinge to the attach points (hinge/control rods) The control rods take the load as compression or tension to its attach point - which in this case is probably a rib - plus - at the hinge/rib attach as shear on the hinge and hinge bearing nestled in wing attach.
This is where you need the drawings. The loads from the Hinge are transferred to the wing as both in plane and normal shear. The normal component of shear is normal also to the spar and normal to the torque box it is introduced to. The in plane shear at the hinge attach is taken "in Plane" to say the Rib where the cross section of the rib is stested for tension load stress.
The Normal forces are then applied to the wing to look for deflections in the torque box corners where the hinges are attached to the wing. The leading edge of the wing is tested for changes to AoA and subsequent aero load distribution for additional normal force introduction in outer sections of the torque box (IMO - This is where the FW 190 ran into trouble with zero LE twist at outer 20% semi span)
Once at the wing root the torsional loads must be transferred from the leading edge box to the rear spar bolt via rib 1. This Rib could have been bolted directly to the fuselage however the load is transferred via the rear spar.
This approach is a distinct possibility - but can't be verified unless and until you size the Rib and look at the engineering analysis. Having said that, you can't have your analogy in which you 'cut' from trailing edge to spar at 24% chord. You would be killing the shear panel from 25% to say 75% (wherever aft spar/flap hinge is located) which in turn would destroy any useful structure to keep the rib "In plane" as a useful stiff member to beam out the loads from the front spar.
The most significant part of the rear spar are those few inches between rib 1 and fuselage attachment. The trailing edge spar outboard of that area is of little significance I assert.
From my perspective it seems very useful to provide continuity for a torque box connecting to fuselage interface rib from chord .25 to ~ .75
Supermarine seems to have been cautious in penetrating the leading edge of the spitfire and of course the holes were not that big or had elaborate engineering around them, such as the plug leading edge nacelles that were plugged and left in place when the 0.5 Brownings or Hispano weren't fitted.
I accept that the rear spar bolt is critical in providing torsional stiffness; however I see the rear spar of almost total insignificance. It could have been eliminated.
Firstly there is almost zero torsional load on the spitfire wing when flying statically since the centre of lift is where the spar is at around 25% of chord.
The Center of Pressure is different from the Aerodynamic Center. The Center of Pressure is a point on the airfoil chord where the moments created by a Normal force acting on the airfoil and an in plane Shear force act on the airfoil - At this point the aerodynamic moment is Zero. It is real, and it is the Centroid of the Lift/pressure distribution - and it is Not at 25% chord.
The Aero Center is at a point (25% Chord) where the moment about the AC (CMac) will not change as AoA changes and is primarily used in Stability and Control analysis.
The CMac is a Calculated Moment to accomodate the transfer of the Forces at the CP to the AC so that AoA effects may be removed from the static stability analytics..CMle would have a different value, for example, if the Lift Forces were translated at the Leading Edge instead of 25% chord.
A casual look through various Aero books will help distinguish the two from each other. i.e page 20 "Fundamentals of Aerodynamics" John D Anderson, McGraw Hill.
Minor torsional loads would come from pitching moment as AOA varied; however aileron torsional loads would be more.
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They also didn't fly in excess of 300 mph and weren't capable of pulling 4 Gs...... Why not- Early WWI aircraft didn't have a rear spar too...
What you see is immaterial, since it wasn't eliminated, therefore it was still there, therefore the Spitfire wing had two spars, not one..I accept that the rear spar bolt is critical in providing torsional stiffness; however I see the rear spar of almost total insignificance. It could have been eliminated..
Obviously your knowledge of the Spitfire falls way below what's needed. The Mk.I/IIa/Va leading edge had tubes fixed to the mainspar and extreme leading edge, through which the .303" Brownings passed, and which stiffened the whole assembly.Supermarine seems to have been cautious in penetrating the leading edge of the spitfire and of course the holes were not that big or had elaborate engineering around them, such as the plug leading edge nacelles that were plugged and left in place when the 0.5 Brownings or Hispano weren't fitted
On what part of the aircraft? Upper wing? The fabric alone would not be able to take high speed loads and this is like comparing apples and oranges.Max g-load Fokker DR1=8,3
What you see is immaterial, since it wasn't eliminated, therefore it was still there, therefore the Spitfire wing had two spars, not one..
Obviously your knowledge of the Spitfire falls way below what's needed. The Mk.I/IIa/Va leading edge had tubes fixed to the mainspar and extreme leading edge, through which the .303" Brownings passed, and which stiffened the whole assembly.
For the cannon, there was a substantial casting fitted to the leading edge, which was strong enough for the eccentrics (used for setting the convergence angles of the cannon, and .5" Browning - when fitted) to be fitted inside, and which held the guns rigidly in place.
The casting(s) in question:-