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A structured Reynolds-Averaged Navier-Stokes solver is directly coupled to a Linear Stability Theory (LST) solver to include the effect of laminar-turbulent transition in the flow simulations. The flow-field variables of the flow solver are used to both find streamlines along which transition can be predicted and to provide the LST code with the required boundary layer profiles. Instabilities included in the analysis are of the Tollmien-Schlichting (TS) and cross-flow (CF) nature relevant to high Reynolds number flows in low turbulence environments. The coupling is fully automated and can therefore be used efficiently in the analysis and design of geometries with external flows. TU Braunschweig Sickle Wing with span-wise varying cross-flow and the Natural Laminar Flow version of the Common Research Model (CRM-NLF) are simulated under various conditions. Applications to these relevant three-dimensional test cases showcase the capability of the method to model the real flow physics. Advantages and challenges of the approach with regard to future design endeavors are discussed.

Recent developments in the SBP-SAT method have made available high-order interpolation operators (Mattsson and Carpenter, SIAM J Sci Comput 32(4):2298–2320, 2010). Such operators allow the coupling of different SBP methods across nonconforming interfaces of multiblock grids while retaining the three fundamental properties of the SBP-SAT method: strict stability, accuracy, and conservation. As these interpolation operators allow a more flexible computational mesh, they are appealing for complex geometries. Moreover, they are well suited for problems involving sliding meshes, like rotor/stator interactions, wind turbines, helicopters, and turbomachinery simulations in general, since sliding interfaces are (almost) always nonconforming. With such applications in mind, this paper presents an accuracy analysis of these interpolation operators when applied to fluid dynamics problems on moving grids. The classical problem of an inviscid vortex transported by a uniform flow is analyzed: the flow is governed by the unsteady Euler equations and the vortex crosses a sliding interface. Furthermore, preliminary studies on a rotor/stator interaction are also presented.

Airplane wings can suffer from flow separation, which greatly decreases their aerodynamic per-formance. The flow separates due to the bound-ary layer possessing insufficient momentum to engage the adverse pressure gradient along the airfoil surface. Flow separation control actively influences the flow such that flow separation is delayed and airfoil performance is improved. In this research flow separation control is ap-plied on a 2D wing with a NACA0018 section with a 0.165 m chord using tangentially direct-ed synthetic jets. The actuators are located in-side the wing and the jet exits from a slot in the upper wing surface. The synthetic jet inhales the low momentum air in the boundary layer during instroke and during the outstroke the air adds momentum to the boundary layer. The actuator, with a piezo-electric diaphragm, has a slot width of 0.25 mm. With this design jet velocities up to 65 m/s have been achieved at an optimum actuation frequency of 900 Hz. A spanwise row of ten actuators is placed inside the wing, such that the slots cover 66% of the wing's span. During wind tunnel experiments forces have been measured using a balance. The tests have been performed at a fixed free stream velocity of 25 m/s (Rec = 2.73×105) and for various actuation frequencies and jet velocities. It is shown that for given actuation frequency a higher jet velocity results in a higher maximum lift coefficient and a corresponding higher stall angle. However, for the performance of the syn-thetic jet actuation, actuation frequency proves to be of greater importance than jet velocity. The best actuation frequency in combination with the maximum jet velocity possible with the present actuator is a dimensionless frequency F+ of 5.9 (1300 Hz) and a momentum coefficient cμ of 0.0014 (maximum jet velocity 32.9 m/s and Velocity Ratio of 1.32). Using these actuation parameters the lift coefficient is increased by 12% and the stall angle by 22%.

This paper, investigates the feasibility of performing a robust numerical simulation of the dynamics of vortex cavitation. An equilibrium cavitation model is employed, which assumes local thermodynamic and mechanical equilibrium in the two-phase flow region. Furthermore, the phase transition does not depend on empirical constants in this model. The computational method assumes a compressible flow together with appropriate thermodynamic equations of state i.e. Tait's equation for the liquid phase, a perfect gas for the vapor phase, and an equilibrium model for the mixture phase. The unsteady compressible Euler equations are employed using a cell-centered structured multi-block finite volume scheme. The viscous terms are not taken into account, because the numerical issues encountered are typically caused by (the discretization of) the inviscid equations. The 3D vortex cavitation simulations using a MUSCL scheme on an Arndt's elliptic hydrofoil are presented. The results show that the common MUSCL schemes are not well-suited for such complex simulations. For an accurate representation of the cavitating flow, the higher-order accurate Weighted Essentially Non-Oscillatory (WENO) schemes are considered. In order to avoid negative density and/or internal energy, the scheme must fulfill the positivity-preserving property. The existing positivity-preserving approaches use limiter functions to preserve the monotonicity of the solution near the discontinuities. It is well-known that the limiter functions introduce considerable amount of artificial dissipation, which is highly undesirable in vortex cavitation simulations. We propose a non-limiting (no slope/flux limiters) positivity-preserving strategy for WENO schemes, which is rather simple and computationally low-cost. Some 1D test case simulations with or without cav-itation are presented to assess the developed positivity-preserving WENO aproach.

In this study a computational method is presented which simulates the presence of a liquid layer on an airfoil and its effect on splashing of Supercooled Large Droplets (SLD). The thin liquid film is expected to have a significant influence on the impact behaviour of SLD. It will arise when the impacting droplets freeze only partially and leave behind a layer of runback water on top of the ice layer. The liquid film is modelled using the wall shear stress and by assuming a linear velocity profile within the water layer. The shear stress is calculated by coupling an integral boundary-layer method to a potential flow method. The SLD splashing model is extended with a deposition model that accounts for impact on a liquid film and includes the solidification time of the droplets. This solidification time is obtained using multiple approaches which are based on either planar solidification or dendritic solidification. Planar solidification is controlled by diffusion and based on the Stefan problem for heat conduction. Dendritic solidification is more rapid and mostly governed by kinetics. The comparison of the catching efficiency with experimental results for a NACA-23012 airfoil shows a significant improvement employing the new deposition model. Also, good agreement is found with the experimental results for the ice accretion on a NACA-0012 airfoil.

Compared to conventional icing additional droplet phenomena have to be accounted for in icing caused by supercooled large droplets (SLD) such as splashing, rebound, breakup and deformation. In this study the effect of the presence of a thin liquid film of water on the surface has been investigated. This liquid layer can arise when SLD droplets freeze only partially following impact on the airfoil. The effect of the liquid film is simulated by using the wall shear stress and by assuming a linear velocity profile in the liquid layer. The shear stress is calculated by coupling an integral boundary-layer method to a potential ow method. An improved splashing model has been implemented in the existing com-putational method. This splashing model consists of a deposition model that accounts for splashing during impact of droplets on a liquid layer. In an extension to this model different solidification models have been investigated to estimate the time of solidification of a liquid splat produced on the surface after impact. One is a planar solidification model which is described by the Stefan problem for heat conduction and which is mostly controlled by diffusion. The second model is based on dendritic solidification, which is rapid and gov-erned by kinetics. The results of the deposition model on SLD ice accretion are compared with data from experiments on a NACA-23012 airfoil and on a NACA-0012 airfoil. Good agreement is found.

Numerical simulations of active flow control have been carried out for the flow around the NACA0018 profile for Mach = 0.15, Re = 2× 106, a = 15o using the Unsteady Reynolds Averaged Navier-Stokes (URANS) equations. Two types of flow control, zero-net-mass jets (synthetic jets) and continuously blowing jets, have been considered to delay the onset of separation. The synthetic jets have been applied to the 2D situation, i.e. infinitely long slits in the spanwise, for which the angle between the jet and surface normal has been varied to study the effect on the separation. For the continuously blowing jets the effect of 3D mixing is taken into account and an optimization of several jet parameters has been carried out to obtain the best result possible.