Unsteady CFD for Transonic Compressible Flows in application to Shock-Impulsive Helicopter Noise
2-d Benchmark: Mach reflection over a wedge
The wedge corner is located at 0.25 from the left boundary and the angle is 46 degrees. The incident shock wave is propagating at the Mach number 2 from left to right. Boundary conditions: solid wall on the wedge surface and fixed values after the shock wave at the inflow boundary. The grid is uniform and rectangular (256x256) with the diagonal positioned on the wedge slope. Density contours:
First order Roe scheme (51kb) and Second order Roe scheme (34kb)2-d Airfoil in a wind tunnel with oscillating downstream pressure
NACA 0012 aerofoil at zero incidence is placed in a wind tunnel. Rotating valve changes periodically the outlet area (outlet pressure). Numerical simulation is performed with the second order Roe scheme with TVD MUSCL flux splitting with MinMod limiter. Boundary conditions: solid wall on the lower and upper boundary, where more accurate fluxes on the boundary were obtained by extrapolating pressure at fictitious, 'reflective' points; specifying stagnation pressure and enthalpy for the inflow and imposing pressure at the outflow boundary. The rest of the variables is extrapolated from interior. There is a simple H-grid with (50x70) on the aerofoil.
Full domain view: Mach number contours (209kb) Density contours (209kb)2-d helicopter blade section in forward flight-like motion
Accelerating/decelerating airfoil in a free stream with varying pitching: a(t)=a3-d Calculation of helicopter blade motion in forward flight
The model rigid blade rotates around rotor hub, pitches around its axis, cones about the rotation plane that is tilted towards horizontal. There is a rectangular linearly twisted rotor blade with NACA 0012 aerofoil for each section used. The computational mesh consists of a number of O-grids stacked in span-wise direction with wrapping around the blade tip (99x30x36). All external boundary conditions including the hub plane are non-reflecting characteristic-type. Animation of one blade revolution corresponding to advancing blade tip Mach number of 0.8 blade motion (89kb) and blade surface pressure distribution (698kb).A similar calculation is conducted for a typical helicopter blade geometry in a forward flight experiment with advancing tip Mach number 0.88. The grid is (127x30x36). For best non-reflecting results the non-reflecting characterstic-type boundary conditions have been used together with a sponge zone near the outer boundaries. To eradicate the effect of grid stretching on the resulting acoustic field a numerical perturbation approach about the base flow has been used together with the main shock-capturing solver. The evolution of shock wave region as the blade is going through different azimuths: shock motion(981kb)
The evolution of pressure iso-surfaces around the blade : isobars(2660kb)
Created by
Sergey Karabasov, 2002