With the increased interaction between the aerodynamic, structures, and controls disciplines in advanced aircraft, it is increasingly important to accurately predict the aerodynamic forces due to control surface deflection. The aerodynamic behavior of oscillating control surfaces has historically been difficult to predict analytically. The Doublet-Lattice Method (DLM)1 has been the de-facto standard unsteady aerodynamic analysis tool in the subsonic regime for decades, but since it is based on the linear potential equation, it has had limited success in computing control surface aerodynamics. In practice, the aerodynamic results from the DLM (or a similar process) were modified using (sometimes large) correction factors to improve correlation with test data. 

These correction factors are typically based on steady wind tunnel data, and are applied to steady and unsteady conditions alike. This is undesirable because the steady corrections will not necessarily be representative of the unsteady flow, and cannot correct the unsteady phase angle. 

With the advent of low-cost, high-power computers, the use of unsteady Computational Fluid Dynamics (CFD) for the calculation of unsteady control surface aerodynamics has recently become more attractive. In this study, CFL3D.AE-BA2 was used to compute unsteady Navier-Stokes solutions of oscillatory control surface motion. This model includes not only the effects of compressibility and transonic shocks, but also accounts for viscous effects.

Correlation of CFD to test data for the BACT model was accomplished by Schuster, et al.3 with a ENS3DAE Navier-Stokes analysis. In the investigation discussed in this paper, the first step was to similarly validate CFL3D. The study performed by Schuster, et al. was used as another source of comparison for the validity of our CFD results.

For cases with small angle of attack and no deflected control surface, linear theory has been shown to be valid outside the transonic regime. When analysis is required for cases with an oscillating control surface, or for cases in the transonic regime, the validity of linear theory breaks down. Flow characteristics are introduced which cannot be predicted by linear theory. A classic example of such flow characteristics occurs in the shock wave present for flows with a supersonic pocket at the surface of the wing. There will be a difference in the results obtained from linear theory and those obtained from test and CFD in the region upstream of the transition back to subsonic flow. Test data and CFD should predict an increase in phase lag in the unsteady pressure results. This behavior will be seen as near zero magnitude for cases in which the supersonic pocket has even higher velocity. This behavior will not be predicted by linear theory, and it is of the interest of this study to see at what Mach number and frequency of control surface oscillation this poor correlation begins to occur.