Variable Sweep Wings and Aerodynamic Characteristics

Up to now, the airplane was considered to be in motion at subsonic speeds. Drag was composed of three main components: skin-friction drag, pressure drag, and induced drag (or drag due to lift). At transonic and supersonic speeds there is a substantial increase in the total drag of the airplane due to fundamental changes in the pressure distribution, due to the wave drag. As seen in the previous section the sweep delays the compressibility effects through the augmentation of the Divergence Mach number, and reduces the wave drag through the use of thinner airfoils.
There is another reason that determine the use of swept wing in the transonic regime and it’s associated with the so-called Whitcomb Transonic Area Rules .
In the early 1950's, Richard T. Whitcomb of the NACA Langley Memorial Aeronautical Laboratory first experimentally demonstrated an aerodynamic principle that has had a profound and far-reaching effect on the entire process of airplane configuration synthesis. According to Donlan "the basic tenet of the area rule states that the wave drag of an airplane configuration depends primarily on the longitudinal distribution of the total cross-sectional area. This concept results in the proposition that the wave drag of a simple equivalent body of revolution (that is, a body having the same longitudinal distribution of total cross-sectional area) would be the same as that of the more complex wing-body arrangement.
Here follows the physic explanation of the transonic theory:
“The chordwise pressure distribution on a plane, finite span, swept back wing varies across the span such that the maximum local velocity is reached much further forward the tip compared to the basic infinite wing distribution. These distorsions at the root and at the tip determine the flow pattern. On a swept wing with a square cut tip, a shock wave first occurs near the tip. This initial tip shock is relatively short in extent. As the Mach number is increased, a rear shock is developed which dominates the flow. Because of the variation of the component of velocity normal to the wing leading edge, streamlines over the wing surface will tend to be curved. However, the flow at the wing root is constrained to follow the fuselage. Therefore, for a straight fuselage, a set of compression waves originate in the wing root region to turn the flow parallel to the fuselage. The compression waves propagate across the wingspan and ultimately may coalesce near the tip to form the rear shock. Thus, it’s clear that the interaction between the fuselage and the swept wing has a significant effect on the transonic flow over the wing.A properly shaped waistlike indentation on the fuselage can be designed to produce a velocity distribution on the wing near the root similar to that on the corresponding infinite span wing”.