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A numerical solution for ‘running wet’ aircraft anti‐icing systems is developed. The model includes breakup of the water film, which exists in regions of direct impingement, into individual rivulets. The wetness factor distribution resulting from the film breakup and rivulet configuration on the surface are predicted in the numerical solution procedure. The solid wall is modelled as a multi‐layer structure and the anti‐icing system used is of the thermal type utilizing hot air and/or electrical heating elements embedded within the layers. Details of the calculation procedure and the methods used are presented.

Responsible for lift generation, the helicopter rotor is an essential component to protect against ice accretion. As rotorcraft present a smaller wing cross-section and a lower available onboard power compared to aircraft, electro-thermal heating pads are favored as they conform to the blades’ slender profile and limited volume. Their optimization is carried out here taking into account, for the first time, the highly three-dimensional (3D) nature of the flow and ice accretion, in contrast to the current state-of-the-art that is limited to two-dimensional (2D) airfoils.

Conjugate heat transfer simulation results are provided by the truly 3D finite element Navier–Stokes analysis package-ICE code, embedded in a proprietary rotorcraft simulation toolkit, with reduced-order modeling providing a time-efficient evaluation of the objective and constraint functions at every iteration. The proposed methodology optimizes heating pads extent and power usage and is versatile enough to address in a computationally efficient manner a wide variety of optimization formulations.

Low-error reduced-order modeling strategies are introduced to make the tackling of complex 3D geometries feasible in todays’ computers, with the developed framework applied to four problem formulations, demonstrating marked reductions to power consumption along with improved aerodynamics.

The present paper proposes a 3D framework for the optimization of electro-thermal rotorcraft ice protection systems, in hover and forward flight. The current state-of-the-art is limited to 2D airfoils.

This paper aims to present an approach based on electrical impedance tomography spectroscopy (EITS) for the determination of water and ice fraction in low-power applications such as autarkic wireless sensors, which require a low computational complexity reconstruction approach and a low number of electrodes. This paper also investigates how the electrode design can affect the reconstruction results in tomography.

EITS is performed by using a non-iterative method called optimal first order approximation. In addition to that, a planar electrode geometry is used instead of the traditional circular electrode geometry. Such a structure allows the system to identify materials placed on the region above the sensor, which do not need to be confined in a pipe. For the optimization, the mean squared error (MSE) between the reference images and the obtained reconstructed images was calculated.

The authors demonstrate that even with a low number of four electrodes and a low complexity reconstruction algorithm, a reasonable reconstruction of water and ice fractions is possible. Furthermore, it is shown that an optimal distribution of the sensor electrodes can help to reduce the MSE without any costs in terms of computational complexity or power consumption.

This paper shows through simulations that the reconstruction of ice and water mixtures is possible and that the electrode design is a topic of great importance, as they can significantly affect the reconstruction results.

This paper aims to investigate the anti-icing performance of the nanosecond dielectric barrier discharge (NSDBD) plasma actuator.

With the Lagrangian approach and the Messinger model, two different ice shapes known as rime and glaze icing are predicted. The air heating in the boundary layer over a flat plate has been simulated using a phenomenological model of the NSDBD plasma. The NSDBD plasma actuators are planted in the leading edge anti-icing area of NACA0012 airfoil. Combining the unsteady Reynolds-averaged Navier–Stokes equations and the phenomenological model, the flow field around the airfoil is simulated and the effects of the peak voltage, the pulse repetition frequency and the direction arrangement of the NSDBD on anti-icing performance are numerically investigated, respectively.

The agreement between the numerical results and the experimental data indicates that the present method is accurate. The results show that there is hot air covering the anti-icing area. The increase of the peak voltage and pulse frequency improves the anti-icing performance, and the direction arrangement of NSDBD also influences the anti-icing performance.

A numerical strategy is developed combining the icing algorithm with the phenomenological model. The effects of three parameters of NSDBD on anti-icing performance are discussed. The predicted results show that the anti-icing method is effective and may be helpful for the design of the anti-icing system of the unmanned aerial vehicle.

This paper aims to propose a method of the safety boundary protection for unmanned aerial vehicles (UAVs) in the icing conditions.

Forty icing conditions were sampled in the continuous maximum icing conditions in the Appendix C of the Federal Aviation Regulation Part 25. Icing numerical simulations were carried out for the 40 samples and the anti-icing thermal load distribution in full evaporation mode were obtained. Based on the obtained anti-icing thermal load distribution, the surrogated model of the anti-icing thermal load distribution was established with proper orthogonal decomposition and Kriging interpolation. The weather research and forecasting (WRF) model was used for meteorological simulations to obtain the icing meteorological conditions in the target area. With the obtained icing conditions and surrogated model, the anti-icing thermal load distribution in the target area and the variation with time can be determined. According to the energy supply of the UAVs, the graded safety boundaries can be obtained.

The surrogated model can predict the effects of five factors, such as temperature, velocity, pressure, median volume diameter (MVD) and liquid water content (LWC), on the anti-icing thermal load quickly and accurately. The simulated results of the WRF mode agree well with the observed results. The method can obtain the graded safety boundaries.

The method has a reference significant for the safety of the UAVs with the limited energy supply in the icing conditions.

This study aims to couple two codes, one able to perform icing simulations and another one capable to simulate the performance of an electrothermal anti-icing system in an integrated fashion.

The classical tool chain of icing simulation (aerodynamics, water catch and impact, mass and energy surface balance) is coupled to the thermal analysis through the surface substrate and the ice thickness. In the present approach, the ice protection simulation is not decoupled from the ice accretion simulation, but a single computational workflow is considered.

A fast approach to simulate advanced anti-icing systems is found in this study.

This study shows the validation of present procedure against literature data, both experimental and numerical.

FOR a number of years now it has been evident that a successor to the well‐tried Vickers Viscount and Convoir 240/340/440 series was required. However, the big problem was to design an aircraft such that its economics and passengerappealweresub‐stantially better than the machines it would ultimately replace. Other important factors which had to be con‐sidered were improved reliability, easier and cheaper maintenance, higher standards of safety and means of reducing ramp times. Furthermore, the difficult choice of passenger capacity and cruising speed had to be made. Probably the easiest decision was to employ the twin‐engine configuration with the power plants placed in the now familiar rear position, one on cither side of the fuselage.

FUEL SYSTEM GENERALTHE standard version or the F.28 has a conventional two‐tank fuel system with an integral tank in each outer wing section contained by the wing torsion box with a total capacity of 2,170 Imp. gal. or 17,200 lb. of fuel. The centre wing torsion‐box pro‐vides space for additional bladder type tanks with a capacity varying from 312 to 700 Imp. gal. or 2,460 to 5,500 lb. as desired.

THERMAL protection from the effects of ice built up can be achieved by means of anti‐icing or de icing. Most aircraft structures are de‐iced since this consumes considerably less power, but in some cases it is essential that the surface be anti‐iced, (e.g. engine intakes, where ice ingestion to the engine is unacceptable and any formation of ice must therefore be prevented).