Transonic Swept Wing Research Articles

Shock control on a swept wing

Shock and boundary layer control by contour bumps and local boundary layer suction have been investigated experimentally and numerically on a transonic swept wing. Additional 2-D numerical investigations were performed for the airfoil, corresponding to the wing. The investigations were primarily stimulated by the question concerned with the influence of sweep on the bump effectiveness. This influence has been found to be rather small; the drag reduction by the bump is slightly lower for the swept wing than for the airfoil. A location of the bump in the shock region has shown its effectiveness for reducing shock strength and hence wave drag. A position of the bump downstream of the shock wave has been shown to reduce viscous drag and to postpone buffet-onset to higher lift coefficients. Furthermore, the results indicate that boundary layer suction is a powerful device for drag reduction, but the effectiveness decreases with increasing Reynolds number. Higher effectiveness of suction can be attained, when it is coupled with a contour bump. The parameters height and position (relative to the shock) of the bump, optimized in terms of drag, depend on the shock strength; an influence of the boundary layer thickness upstream of the shock on the optimal bump parameters has not been found. A possibility to control an adaptive bump, mounted on an aircraft wing, is to employ the trailing edge pressure.

Transonic swept wings studied by the lifting-line theory

Transonic swept wings are analyzed as a lifting-line problem under a small-disturbance approximation. Basic concepts and principal results of the asymptotic theory are discussed. The study focuses on straight oblique wings and V-shaped swept wings, of which the local centerline curvature can be equated to zero. The three-dimensional (3-D) perturbation of the nonlinear component flow admits a similarity flow structure but requires that all wing sections are generated from a single airfoil profile; the reduced 2-D problems in this case are solved only once for all span stations. Examples of solutions involving high subcritical and slightly supercritical component flows are demonstrated and compared with surface pressure data from 3-D computer codes based on the full-potential equation (FLO 22). Except in the neighborhood of leading edges, where the small-disturbance assumption breaks down, and in the vicinities of wing tips and the symmetry plane, where neither the theory nor the 3-D codes may claim full validity, reasonable agreement is consistently found. The explicit results from the upwash analysis, along with the similarity flow structure, provides a rational approach to the control of 3-D effects in transonic aerodynamic design studies.

Lifting-line theory of swept wings based on the full potential theory

The asymptotic theory of high-aspect-ratio transonic swept wings of Refs. [2–5] is extended to irrotational compressible flows based on equations of the full-potential theory. The formulation allows for imbedded supercritical flows with shocks; the results are applicable to oblique wings, swept-back as well as swept-forward wings, but restricted to planforms with centerlines composed of straight segments. The free-stream velocity is assumed to remain below the sonic speed. For a wing geometry generated from a single airfoil shape, the reduced system of 2-D equations admits solutions with similarity structures, permitting the problem to be solved only once for all span stations. The application of the theory is expected to lead to a substantial improvement over the earlier analysis of transonic swept wings based on the small-disturbance theory, concerning the accurate determination of inviscid drag and lift. The study provides also a basis for examining the influence of the trailing vortex sheet geometry, which are incompletely accounted for in current 3-D computer programs.

The Linear Instability Due to the Compressible Crossflow on a Swept Wing

The problem of the propagation of three-dimensional laminar instabilities, due to crossflow, in a three-dimensional compressible boundary layer, is examined using linear theory. The theory is applied to the case of a transonic swept wing. It is shown that compressibility has a mild stabilizing effect in the regions where the crossflow is strong. The problem of defining the direction of propagation of the disturbances is discussed.