An inertial particle separator (IPS) mounted before a helicopter engine faces the danger of ice accretion caused by the impingement of super-cooled water droplets. This paper shows the results of a numerical study on the flow and impingement of water droplets in the IPS using the Eulerian–Eulerian method. Cases for different medium volume diameters (MVDs = 5 μm, 10 μm, and 20 µm), liquid water contents (LWCs = 0.5 g/m³, 2 g/m³, and 4 g/m³), and incoming velocities (U₀s = 40 m/s, 65 m/s, and 90 m/s) are examined to study their effects on the flow and impingement of water droplets in the IPS. The results indicate that the IPS can reduce the water that enters the core flow path by trapping the water on the inner walls and separating it into the scavenge flow path. The ratio between the flow rate of water at the outlet of the core flow path and that at the inlet (${\dot{m}}_c/{\dot{m}}_{inlet}$ ) decreases with an increased MVD or incoming velocity, which, however, does not change with the LWC. The results also show that the water collection coefficient (β) on the hub and shroud walls increases with a greater MVD, while the β on the splitter wall reaches its maximum at an MVD of 10 µm in the studied cases. The LWC does not have any effect on β for any of the IPS walls, while β increases on all IPS surfaces with the incoming velocity.

Aircraft icing seriously threatens flight safety. This paper describes a modified shallow-water icing thermodynamic model that is applicable to unstructured grids and considers the effects of changes in water physical parameters and sublimation on icing. An icing calculation method enables the automatic determination of whether freezing occurs and the type of icing. An autonomous icing calculation program is developed to extract the geometric coordinates and mesh information of the icing surface and combine this information with the multiphase flow field of air-supercooled water droplets. The icing equations in the Godnov format are discretized to produce a large-scale sparse matrix that is solved iteratively using the biconjugate gradient stabilized method. The output includes parameter distributions for the ice thickness, water film thickness, and equilibrium temperature. Taking the National Advisory Committee for Aeronautic 0012 airfoil as the study object, the results of icing simulations are compared with experimental data from an ice wind tunnel and the ice shapes calculated by the FENSAP-ICE software. The results are found to be in good agreement with the experimental data, and the ice height errors at stagnation points are less than 15%. In the case of rime ice, single-horned ice shapes are simulated for both conventional and large supercooled droplets. For mixed ice and glaze ice, double-horned ice shapes are simulated for both droplet conditions. FENSAP-ICE fails to simulate the ice horns and produces large errors in ice thickness and ice range.

This paper proposes a three-dimensional model to simulate ice accretion on the surface of an aircraft. The model is developed using a prism grid cell, which enables iterations with an unstructured mesh. The model assumes that the impinged water droplets on the surface can form a thin water film, and ice accretion is caused by the flow and solidification of the film. The model is developed by analyzing the conservations of mass, momentum, and energy for the thin water film on the surface. The mass and energy conservation equations of the water film are discretized on the prism grid cell, while the momentum conservation equation, which is simplified using lubrication theory, is integrated along the surface normal direction to obtain an analytical solution for the film velocity. The thicknesses of the water film and ice layer are calculated by iterating the discretized mass and energy conservation equations with the analytical solution of the film velocity. The model is verified by comparing its results with published experimental and numerical data. The results show that the model provides accurate predictions of ice accretion under different conditions, even when using an unstructured grid.

When an aero engine operates under icing conditions, the windward surfaces of its intake components, such as the zero-stage guide vane, are prone to icing, posing a potential flight safety hazard. In this study, we establish a computational model for hot-air anti-icing of a full-scale zero-stage guide vane based on CFX and FENSAP-ICE software and conduct a numerical study on three-dimensional coupled heat transfer for anti-icing and the improvement of anti-icing structures. The computational results demonstrate that under anti-icing conditions, the lowest temperature is observed at the root of the leading edge of the guide vane. Icing predominantly occurs in the water droplets impinging on the area downstream of the stagnation point on the guide vane and in proximity to the root of the suction surface, with ice reaching a maximum thickness of ∼7.21 mm. When the hot-air inlet is shifted toward the leading edge and a deflector is introduced, the anti-icing surface temperature is elevated, the icing area is slightly reduced, and the total ice mass is decreased by ∼14.7% compared to the original model. Furthermore, after designing opening holes near the root of the suction surface to allow the hot air to flow out, the temperature on the guide vane’s surface is significantly increased, and the temperature gradient is reduced. The icing area is markedly decreased, with the maximum ice thickness reduced by ∼3.3 mm, and the total ice mass further decreased by about 33.1%.