Editing Swept wing (section)

===Subsonic and transonic flight===
[[File:Jak 25.svg|thumb|[[Yakovlev Yak-25]] swept wing ]]
[[File:swept wing w transonic shock.svg|thumb|Shows a swept wing in transonic flow with the position of a shock wave(red line). This line is a line of constant pressure (isobar) since shock waves cannot exist across isobars and for a well-designed wing coincides with a constant percent chord<ref>Fundamentals Of Flight, Second Edition, Richard S.Shevell{{ISBN|0 13 339060 8}}, p.200</ref> as shown. The triangles show that only part of the streamwise incident airflow is responsible for producing lift or causing shock waves (i.e. that part shown by the arrow perpendicular to the red isobar). Its length behind the shock is shorter signifying that the flow has slowed down in going through the shock.]]

Shock waves can form on some parts of an aircraft moving at less than the speed of sound. Low-pressure regions around an aircraft cause the flow to accelerate, and at transonic speeds this local acceleration can exceed Mach 1. Localized supersonic flow must return to the freestream conditions around the rest of the aircraft, and as the flow enters an adverse pressure gradient in the aft section of the wing, a discontinuity emerges in the form of a shock wave as the air is forced to rapidly slow and return to ambient pressure.

At the point where the density drops, the local speed of sound correspondingly drops and a shock wave can form. This is why in conventional wings, shock waves form first ''after'' the maximum Thickness/Chord and why all airliners designed for cruising in the transonic range (above M0.8) have supercritical wings that are flatter on top, resulting in minimized angular change of flow to upper surface air. The angular change to the air that is normally part of lift generation is decreased and this lift reduction is compensated for by deeper curved lower surfaces accompanied by a reflex curve at the trailing edge. This results in a much weaker shock wave towards the rear of the upper wing surface and a corresponding ''increase'' in critical mach number.

Shock waves require energy to form. This energy is taken out of the aircraft, which has to supply extra [[thrust]] to make up for this energy loss. Thus the shocks are seen as a form of [[drag (physics)|drag]]. Since the shocks form when the local air velocity reaches supersonic speeds, there is a certain "[[critical mach]]" speed where sonic flow first appears on the wing. There is a following point called the [[drag divergence mach number]] where the effect of the drag from the shocks becomes noticeable.  This is normally when the shocks start generating over the wing, which on most aircraft is the largest continually curved surface, and therefore the largest contributor to this effect.

Sweeping the wing has the effect of reducing the curvature of the body as seen from the airflow, by the cosine of the angle of sweep. For instance, a wing with a 45 degree sweep will see a reduction in effective curvature to about 70% of its straight-wing value. This has the effect of increasing the critical Mach by 30%. When applied to large areas of the aircraft, like the wings and [[empennage]], this allows the aircraft to reach speeds closer to Mach 1.

One limiting factor in swept wing design is the so-called "middle effect". If a swept wing is continuous - an [[Oblique wing|oblique swept wing]] - the pressure isobars will be swept at a continuous angle from tip to tip. However, if the left and right halves are swept back equally, as is common practice, the pressure isobars on the left wing in theory will meet the pressure isobars of the right wing on the centerline at a large angle.  As the isobars cannot meet in such a fashion,{{why|date=July 2022}} they will tend to curve on each side as they near the centerline, so that the isobars cross the centerline at right angles to the centerline. This causes an "unsweeping" of the isobars in the wing root region.  To combat this unsweeping, German aerodynamicist [[Dietrich Küchemann]] proposed and had tested a local indentation of the fuselage above and below the wing root. This proved to not be very effective.<ref name="GerDev">Meier, Hans-Ulrich, editor ''German Development of the Swept Wing 1935–1945'', AIAA Library of Flight, 2010. Originally published in German as ''Die deutsche Luftahrt Die Pfeilflügelentwicklung in Deutschland bis 1945'', Bernard & Graefe Verlag, 2006.</ref> During the development of the [[Douglas DC-8]] airliner, uncambered airfoils were used in the wing root area to combat the unsweeping.<ref>Shevell, Richard, "Aerodynamic Design Features", DC-8 design summary, February 22, 1957.</ref><ref>Dunn, Orville R., "Flight Characteristics of the DC-8", SAE paper 237A, presented at the SAE National Aeronautic Meeting, Los Angeles California, October 1960.</ref>