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Session: Low/high order methods 1
Session starts: Wednesday 19 June, 13:30
Presentation starts: 14:30
Room: Room 1.3

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Lipeng Zhu ()
Changchuan Xie ()
Yang Meng ()

Abstract:
The objective of this paper is to develop an aerodynamic model suitable for aeroelastic analysis with low computational cost and sufficient fidelity. The physics-based reduce order model is based on the unsteady inviscid Panel Method (PM), selected for its low computation time. Viscous effects are modeled with two-dimensional unsteady high-fidelity boundary layer calculations at various sections along the span and incorporated as an effective shape boundary condition correction inside the PM. The viscous sectional data are calculated with two-dimensional differential boundary layer equations to allow viscous effects to be included for a more accurate maximum lift coefficient and spanload evaluations. These viscous corrections are coupled through a modified displacement thickness distribution coupling method for 2D boundary layer sectional data. Predicting the flowfield by solutions based on inviscid-flow theory is usually adequate as long as the viscous effects are negligible. A boundary layer that forms on the surface causes the irrotational flow outside it to be on a surface displaced into the fluid by a distance equal to the displacement thickness, which represents the deficiency of mass within the boundary layer. Thus, a new boundary for the inviscid flow, taking the boundary-layer effects into consideration, can be formed by adding to the body surface. The new surface is called the displacement surface and, if its deviation from the original surface is not negligible, the inviscid flow solutions can be improved by incorporating viscous effects into the inviscid flow equations. For a given wing geometry and freest ream flow conditions, the inviscid velocity distribution is first obtained with the three-dimensional panel method, and the boundary layer equations are solved along the streamline. The fidelity of the method is verified against 3D RANS flow solver solutions on a high aspect ratio wing. The overall results show impressive precision of the 3D PM/2D BL approach compared to 3D RANS solutions and in compute times in the order of minutes on a standard desktop computer. The steady and unsteady analysis results of NACA 0012 airfoil are shown as follows.