Shock migration on an oscillating straked delta wing using an unsteady Euler solver

Many investigations, including flight tests, wind tunnel tests, and computational methods have attempted to determine the drivers of limit cycle oscillations (LCO) on the F-16 fighter aircraft. One area of interest has focused on shock-induced trailing-edge separation (SITES) as a driver of limit cycle oscillations. The investigation presented herein aims to expand the understanding of unsteady, nonlinear aerodynamics by studying the flow over a wing undergoing pitch oscillations using Euler computational fluid dynamics with a boundary layer coupling scheme capable of estimating viscous flow effects within the boundary layer. A straked NACA 64A204 delta wing model was used for this analysis. Parameters such as trim angle of attack, pitch amplitude, and Mach number were varied to create 71 unique test cases. Results demonstrate that the same single shock systems at low angles of attack and two shock systems at moderate to high angles of attack that were observed by Cunningham in his wind tunnel tests were also present in this investigation. Patterns of significant shock movement and disappearance were observed within the transonic flow region. Hysteresis of the pitching moment was also found to be present. These findings support the theories that shock movement and SITES may play significant roles in driving LCO.