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If you search Google for span wise flow, the first result under the pictures tab is a image of a wing with the span wise flow at the trailing edge:

enter image description here
(Source)

If you look at some other photos you notice others pictures that show to span wise flow at the leading edge:

enter image description here
(Source)

Is the flow at the leading or trailing edge?

At first I though the flow must be at the trailing edge because this would cause the airflow to point more and more towards the tip of the wing. After a little thought I realized that the air(span wise flow) should be flowing along the leading edge because that is where the air splits up in the first place.

So basically I am asking if the span wise flow is flowing along the trailing or leading edge of the wing.

How would this air cause the airflow to gradually point towards the wingtip?

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    $\begingroup$ @mins is right, the decomposition is a theoretical tool that we use to understand the flow. The air does not 'splits up' in to air that goes chordwise and air that goes spanwise. It's the same as when you have a diagonal force, and you split it up in horizontal and vertical forces. The force is still diagonal, but we do the split up to make the evaluation easier. $\endgroup$ – ROIMaison May 31 '17 at 7:43
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The two pictures in the question only show how to decompose the speed vector into the normal and parallel components. They are nowhere near "real" flow vectors.

First I recommend to read this answer so you get an idea how air is accelerated and decelerated when flowing around an airfoil. Note that the low pressure area over the wing will suck in more air. This answer might also be helpful because it looks closer at the boundary layer.

With a swept wing the same happens, but now the flow will not stay at one wing section but show an added sideways component. Again, air is sucked into the low pressure area, but because that area is a little to the side of the direction of motion of the wing, the flow lines will initially move towards the center of the wing on the upper side in case of sweepback.

But that cannot last: The pressure recovery over the rear part of the airfoil reduces this bending, so the flow lines become almost parallel to the direction of motion of the wing. The lower the local pressure, the more the flow lines bend inward. Higher than ambient pressure at the trailing edge or on the lower side of the wing will bend the flow lines outward accordingly.

Flow lines approaching and over a swept wing

Let the red color denote suction: It bends the flow lines (black) towards it, creating a spanwise flow component. The higher the suction (the redder the color), the bigger the bending. The cyan arrows show the flow lines inside the boundary layer. By losing some of the early acceleration to friction but being subject to the same deceleration in the region of the pressure rise, the boundary layer flow shows a stronger tendency towards the tip. To be precise: The acceleration and deceleration acts in a direction perpendicular to lines of the local relative chord. The sketch above exaggerates this bending for the sake of clarity.

Now we must look at the boundary layer to get the full picture: Here the flow speed is reduced by friction, so the result of that initial inbound acceleration is worn away gradually. The result is a twist of the local speed vectors over the thickness of the boundary layer such that the inbound component is gradually reduced the closer you get to the wing's skin. Now the deceleration in the area of pressure recovery will slow down the already decelerated air close to the wing, and again that slowing will happen in a direction perpendicular to the lines of equal chord. If this deceleration is sufficient, it will leave the boundary layer flow with an outward speed component. Such a deceleration brings the flow close to separation so it occurs only at higher angle of attack when the wing approaches stall. Note that only the boundary layer will exhibit a tipwise speed component.

Only when the flow separates, the boundary layer thickens past the separation line (blue line in the sketch above) and the spanwise flow will become significant. Separated flow is characterized by a local forward speed component, and now, past the separation line, you get a layer of air which is left with mostly a spanwise component; outboard in case of a sweptback wing and inboard in case of a forward swept wing.

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  • $\begingroup$ I've read it four times and am still blinking. I wish I still had the books of the professor who could successfully explain concepts to freshly hungover students who had been sampling beers 1-10 in the Belgian Beer Cafe the evening before. $\endgroup$ – Koyovis Jun 3 '17 at 4:09
  • $\begingroup$ @Koyovis: If you could be a little more specific, I could try and improve the answer. $\endgroup$ – Peter Kämpf Jun 3 '17 at 5:06
  • $\begingroup$ Where is the resulting total spanwise flow towards the tip introduced at higher angle of attack? $\endgroup$ – Koyovis Jun 3 '17 at 6:18
  • $\begingroup$ @Koyovis: Only the boundary layer will flow outward. In case of separation, this can be massive, but without separation this is limited to a thickening of the boundary layer on the outer wing. The outer flow will only show a tipwise speed component if local pressure is above ambient pressure. $\endgroup$ – Peter Kämpf Jun 3 '17 at 8:28
  • $\begingroup$ @Koyovis: Is it any better now? $\endgroup$ – Peter Kämpf Jun 4 '17 at 10:20
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Those pictures are valid at the leading edge, and at the trailing edge, and at all points in between on the wing area. But that is only an explanation on why wing sweep delays supersonic shock. In the pictures in the question, the real air flow is the black arrow flow marked Total, and that does not stream toward the tip. There is more to the story.

In the diagrams the black arrow marked Total is perpendicular with the fuselage. If a swept back wing is tilted at a higher angle of attack, the black arrow marked Total will start to deflect towards the wing tips. A partly stalled wing profile looks like this: enter image description here After the separation point there is low pressure. When we look at the top side of a swept back wing, the low pressure line runs along the trailing edge. Consider the red dot at the trailing edge: it sees a region with high pressure to one side (towards the root, coming from the bottom of the wing) and low pressure towards the tip, and air starts flowing towards the tip. So the streamlines look like in this picture. enter image description here So span wise flow is:

  • The Spanwise Component in the pictures in the question

  • An extra component that bends the airstream towards the wing tip.

This causes swept back wings to stall tip first, which causes a nose up pitching moment (which increases the stall) and loss of aileron control. Not desirable effects, and to be limited as much as possible.

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  • $\begingroup$ Not sure what you mean, but have put a smaller comprehension step in the phrasing of the answer and will add more detail to the pictures when I have a chance. I find answers with too much unnecessary detail in the phrasing distracting. $\endgroup$ – Koyovis May 31 '17 at 21:36
  • $\begingroup$ But, why does the total arrow point more and more at the wingtip before the wingtip is stalled at all. $\endgroup$ – Crafterguy Jun 19 '17 at 18:50
  • $\begingroup$ Before flow separation that hardly happens, only in the boundary layer, as @Peter Kämpf explains. At low angles of attack, for understanding purposes, it is probably best to consider 99.9 % of the air flow streaming straight through, like the black arrows in your picture. Spanwise flow is then just a vector component which explains why sweepback works in delaying compressibility effects. Real spanwise flow as in my picture is only consequential at high angles of attack, in explaining the bad stall characteristics of highly swept wings. $\endgroup$ – Koyovis Jun 20 '17 at 2:05

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