Runback Ice Research Articles

Effects of Wind Speed and Heat Flux on De-Icing Characteristics of Wind Turbine Blade Airfoil Surface

The icing on wind turbines reduces their aerodynamic performance and can cause other safety issues. Accordingly, in this paper, the de-icing characteristics of a wind turbine blade airfoil under different conditions are investigated using numerical simulation. The findings indicate that when the de-icing time is 10 s, the peak ice thickness on the leading edge of the airfoil surface decreases from 0.28 mm to 0.068 mm and from 0.77 mm to 0.45 mm at low (5 m/s) and high (15 m/s) wind speeds, respectively. This is due to the fact that the ice melting rate is much greater than the icing rate at low wind speeds, while the icing rate increases at high wind speeds. When the de-icing time is 20 s, ice accretion on the leading edge of the airfoil is completely melted. At a low heat flux (8000 W/m2) and high heat flux (12,000 W/m2), the peak ice thickness decreases by 31.2% and 64.9%, respectively. With an increase in de-icing time and heat flux, the peak thickness of runback ice increases. This is due to an increase in runback ice as a result of more ice melting on the leading edge of the airfoil. The surface temperature in the ice-free area is significantly higher than that in the ice-melting area, due to high thermal resistance in the ice-free area. This study will provide guidance for the thermal distribution and coating layout of a wind turbine blade airfoil to make the anti-/de-icing technology more efficient and energy-saving.

Design and optimization of a modular Electro-Thermal Ice Protection System for a Horizontal Axis Wind Turbine

In this work, an Electro-Thermal Ice Protection System (IPS) design was refined and optimized on the NREL 5 MW reference wind turbine. These improvements involved optimization of the distribution of heat flux in both the stream-wise and span-wise directions to minimize power consumption and avoid run-back ice formation. The design conditions focus on a 3-hour rime ice accretion event. The chosen optimization algorithm is a derivative-free method called Mesh Adaptive Direct Search. his optimization effort results in a 55% reduction in power consumption necessary to maintain a completely clean blade surface. The power and energy losses associated with this new optimized design are compared with those from various ice shapes and with the clean blade conditions established in the initial design phase from previous work. The new optimized design allows to keep the blade clean from ice with much larger energy saving. Indeed, only initially, during the first two hours of ice accretion, the energy losses due to this IPS design are higher than without IPS. Beyond this short initial period, substantial energy savings are achieved, especially considering that after the accretion event, the blade remains ice-free, and there are no more power losses.

Optimisation of a Nacelle Electro-Thermal Ice Protection System for Icing Wind Tunnel Testing

Abstract Aircraft are equipped with ice protection systems (IPS), to avoid, delay or remove ice accretion. Two widely used technologies are the thermo-pneumatic IPS and the electro-thermal IPS (ETIPS). Thermo-pneumatic IPS requires air extraction from the engine negatively affecting its performances. Moreover, in the context of green aviation, aircraft manufacturers are moving towards hybrid or fully electric aircraft requiring all electric on-board systems. In this work, an ETIPS has been designed and optimised to replace the nacelle pneumatic-thermal system. The aim is to minimise the power consumption while assuring limited or null ice formation and that the surface temperature remains between acceptable bounds to avoid material degradation. The design parameters were the length and heat flux of each heater. Runback ice formations and surface temperature were assessed by means of the in-house developed PoliMIce framework. The optimisation was performed using a genetic algorithm, and the constraints were handled through a linear penalty method. The optimal configuration required 33% less power with respect to the previously installed thermo-pneumatic IPS. Furthermore, engine performance is not affected in the case of the ETIPS. This energy saving resulted in an estimated reduction of specific fuel consumption of 3%, when operating the IPS in anti-icing mode.

Icing on Full-Scale Swept Wings in Flight and Icing Wind Tunnel

Inflight ice accretion on aircraft wings has been critically important for development and certification of transport category aircraft. With recent efforts toward certification by analysis, it is becoming increasingly important to characterize ice accretions in icing wind tunnels, the primary source of validation data for higher-fidelity three-dimensional ice shape analysis codes to be used in aircraft certification, in comparison with those in natural icing environments. In this study, a series of natural icing flight tests and icing wind tunnel tests were conducted on a transport category aircraft. The results were compared at similar test conditions on protected and unprotected surfaces of swept wings and empennage. For unprotected surfaces, both tests resulted in scallop ice, with natural icing flight test resulting in significantly smaller ice accretions. For protected surfaces, runback ice was formed on the lower wing surface in both tests, with natural icing flight test resulting in significantly less and sparser runback ice. Differences in scallop ice structures, limitations of icing wind tunnels, and scaling methods are considered the potential contributing factors to the differences. The results indicate a need for further studies on case-to-case comparison of ice accretions in the two test methods focusing on full-scale aircraft conditions.

Design and Development of an Experimental Setup of Electrically Powered Spinning Rotor Blades in Icing Wind Tunnel and Preliminary Testing with Surface Coatings as Hybrid Protection Solution

In order to study ice protection systems for rotating blades, a new experimental setup has been developed at the Anti-Icing Materials International Laboratory (AMIL). This system consists of two small-scale rotating blades in a refrigerated icing wind tunnel where atmospheric icing can be simulated. Power is brought to the blades through a slip ring, through which the signals of the different sensors that are installed on the blades also pass. As demonstrated by the literature review, this new setup will address the need of small-scale wind tunnel testing on electrically powered rotating blades. To test the newly designed apparatus, preliminary experimentation is done on a hybrid ice protection system. Electrothermal protection is combined with different surface coatings to measure the impact of those coatings on the power consumption of the system. In anti-icing mode, the coatings tested did not reduce the power consumption on the system required to prevent ice from accumulating on the leading edge. The coatings however, due to their hydrophobic/superhydrophobic nature, reduced the power required to prevent runback ice accumulation when the leading edge was protected. One of the coatings did not allow any runback accumulation, limiting the power to protect the whole blades to the power required to protect solely the leading edge, resulting in a potential 40% power reduction for the power consumption of the system. In de-icing mode, the results with all the substrates tested showed similar power to achieve ice shedding from the blade. Since the coatings tested have a low icephobicity, it would be interesting to perform additional testing with icephobic coatings. Also, a small unheated zone at the root of the blade prevented complete ice shedding from the blade. A small part of the ice layer was left on the blade after testing, meaning that a cohesive break had to occur within the ice layer, and therefore impacting the results. Improvements to the setup will be done to remedy the situation. Those preliminary testing performed with the newly developed test setup have demonstrated the potential of this new device which will now allow, among other things, to measure heat transfer, force magnitudes, ice nucleation, and thermal equilibrium during ice accretion, with different innovative thermal protection systems (conductive coating, carbon nanotubes, impulse, etc.) as well as mechanical systems. The next step, following the improvements, is to measure forced convection on a thermal ice protection system with and without precipitation and to test mechanical ice protection systems.

Effects of flow parameters on thermal performance of an inner-liner anti-icing system with jets impingement heat transfer

The multiple jets impingement heat transfer is widely applied in the wing anti-icing system. It is challenging to apply the similarity criterion to carry out the anti-icing experiments due to the complex flow and heat transfer behavior. In the present study, the full-scale slat model is used to carry out anti-icing experimental researches in a 2 m × 3 m icing wind tunnel of China Aerodynamics Research and Development Center. The effects of icing parameters Liquid Water Content (LWC) and Median Volume Diameter (MVD) and hot air parameters (mass flow rate and temperature) on the thermal performance of an inner-liner anti-icing system with jets impingement heat transfer are studied. The effects of the experimental parameters are analyzed in detail by combining impingement and evaporation heat transfer mechanisms. The impingement hot air mass flow rate dramatically affects the heat transfer performance of the impingement stagnation region within the range of the experimental parameters. The temperature of impingement hot air and that of wing skin are approximately linear correlated. The experimental results show the effects of LWC and MVD on water film formation and runback ice accretion. The formation of water film is analyzed by an analytical method based on the wing skin temperature difference of dry and wet air conditions.

Numerical simulation of a thermal Ice Protection System including state-of-the-art liquid film model

In this work, a numerical model for the operation of an electro-thermal Ice Protection Systems (IPS) for an airfoil is presented. The present model solves the energy conservation laws and includes a boundary layer model to compute the relevant aerodynamic quantities based on previous works. Additionally, the computation of aerodynamic and water impingement properties is performed by means of open-source and in-house developed software. A state-of-the-art liquid film model based on lubrication theory has been deployed as an alternative to an element-wise mass balance. A simple ice formulation for the prediction of the formation of runback ice has been included. Moreover, a robust evaporation model based on the heat and mass transfer analogy is deployed. Finally, an interpolation scheme for the element-wise enthalpy variation is presented. The results obtained are in good agreement with the experimental data from the open literature for a range of different test cases including different operating modes, environmental and flight conditions.

Effect of ice crystals on run-back ice evolution on a wing model

Effect of ice crystals on run-back ice evolution on a wing model

An experimental study on different plasma actuator layouts for aircraft icing mitigation

An experimental study was conducted to examine the effects of dielectric-barrier-discharge plasma (DBD) actuator layout on the plasma-induced thermal characteristics and evaluate their effectiveness for aircraft icing mitigation. The experimental investigation was performed in the Icing Research Tunnel of Iowa State University (i.e., ISU-IRT) with an airfoil/wing model embedded with an array of DBD plasma actuators around the airfoil leading edge. The plasma actuators were arranged in different layouts (e.g., orientation, number, and width of exposed electrodes) to evaluate their effects on the anti-icing performance under a typical glaze icing condition pertinent to aircraft inflight icing phenomena. While the dynamics ice accretion or anti-icing process over the airfoil surface before and after turning on the plasma actuators was recorded by using a high-resolution imaging system, a high-speed infrared thermal imaging system was used to quantitatively map the temperature distributions over the airfoil surface. The experimental results clearly reveal that, with the same power consumption level, the plasma actuators in streamwise layout would result in higher plasma-induced surface heating and faster runback of the unfrozen water over the airfoil surface, thereby, having a noticeably better anti-icing performance, in comparison to those in spanwise layout. The plasma actuators in streamwise layout were found to not only be able to prevent ice accretion near the airfoil leading edge, but also allow the plasma-induced surface heating to convect further downstream to delay/prevent the runback ice formation near the airfoil trailing edge. A proper combination of the plasma actuation strength and the plasma discharge coverage over the airfoil surface was found to result in the optimum performance for aircraft anti-icing applications.

The influence of ice crystals on the evolution of run-back ice in mixed icing conditions

In the framework of the development of physical and mathematical models of icing in crystalline and mixed conditions and their verification, the effect of ice crystals in the airflow on the evolution of run-back ice on the surface of a wing model for tests in icing facility is numerically studied. The results confirm that the experimentally observed change in the mass of ice deposits with the introduction of crystals into the flow is associated with the absorption of a part of the mass of crystals by the film at low flow rates and with the film splashing at higher speeds.

Evolution of the water film and run-back ice on the surface of a body in plane airflow

A physical-mathematical model of cooling and solidification of a liquid film entrained by air along the heated surface of a streamlined body with given distribution of the heat flux density over the surface is developed. An example of model numerical testing with a set of governing parameters characteristic of experiments at an air cooling rig is presented.

Ice accretion and aerodynamic effects on a multi-element airfoil under SLD icing conditions

The impingement behavior of large water droplets, their interactions with the solid wall and the subsequent ice accretion and aerodynamic effects have become a key issue in in-flight aircraft icing. In this study, ice accretion and aerodynamic effects on a multi-element airfoil were investigated under the recently introduced Appendix O icing envelope. Supercooled large droplet (SLD) dynamics were taken into account by employing a unified computational approach. Ice accretion was simulated using a partial differential equation (PDE) based solver, instead of the commonly used control volume method. The numerical solver of the SLD impingement was built on the droplet deformation and droplet–wall interaction splash models. The unified solvers for clean air, large droplet impingement, ice accretion, and the aerodynamic analysis of ice effects—all of which are based on a single unstructured upwind finite volume framework—were first validated using available experimental data and then applied to investigate ice accretion and the resulting aerodynamic effects on multi-element airfoils for various flight conditions and, in particular, near-freezing SLD icing conditions. Interestingly, two counter-intuitive results were found when comparing the ice accretion and associated aerodynamic degradation for non-SLD and SLD cases. Moreover, considering runback ice was shown to be essential in the design of an ice protection system (IPS) for the multi-element wing.

Temperature and Runback Ice Prediction Method for Three-Dimensional Hot Air Anti-Icing System

A prediction method of surface temperature and runback ice for a three-dimensional hot air anti-icing system was proposed. Computational approach to realize this method was introduced. Both the external and internal flows were separately calculated, results of which were set as boundary conditions of heat conduction computation in airfoil skin. The results of external and internal flow calculations show that the effect of surface temperature on convective heat transfer coefficients and local droplet collection efficiency is negligible and the calculations can be decoupled. The prediction method based on heat flux was used to calculate surface temperature and runback ice results. The results show that, the effects of LWC and Mach number are much more significant than the effect of external flow temperature. The surface temperature at impinging interaction point is more sensitive to the change of external conditions than that at stagnation point. The surface temperature changes significantly with changing Mach number because both the mass rate of droplet and the impact limit are changed.

Performance and durability tests of smart icephobic coatings to reduce ice adhesion

The accretion of ice can cause damage in applications ranging from power lines and shipping decks, to wind turbines and rail infrastructure. In particular on aircraft, it can change aerodynamic characteristics, greatly affecting the flight safety. Commercial aircraft are therefore required to be equipped with de-icing devices, such as heating mats over the wings. The application of icephobic coatings near the leading edge of a wing can in theory reduce the high power requirements of heating mats, which melt ice that forms there. Such coatings are effective in preventing the accretion of runback ice, formed from airborne supercooled droplets, or the water that the heating mats generate as it is sheared back over the wing's upper surface. However, the durability and the practicality of applying them over a large wing surface have been prohibitive factors in deploying this technology so far.Here, we evaluated the ice adhesion strength of four non-conductive coatings and seven thermally conductive coatings by shearing ice samples from coated plates by spinning them in a centrifuge device. The durability of the coating performance was also assessed by repeating the tests, each time regrowing ice samples on the previously-used coatings. Contact angle parameters of each coating were tested for each test to determine influence on ice adhesion strength. The results indicate that contact angle hysteresis is a crucial parameter in determining icephobicity of a coating and hydrophobicity is not necessarily linked to icephobicity.

Experimental study of ice accretion effects on aerodynamic performance of an NACA 23012 airfoil

In this paper, the effects of icing on an NACA 23012 airfoil have been studied. Experiments were applied on the clean airfoil, runback ice, horn ice, and spanwise ridge ice at a Reynolds number of 0.6×106 over angles of attack from −8° to 20°, and then results are compared. Generally, it is found that ice accretion on the airfoil can contribute to formation of a flow separation bubble on the upper surface downstream from the leading edge. In addition, it is made clear that spanwise ridge ice provides the greatest negative effect on the aerodynamic performance of the airfoil. In this case, the stall angle drops about 10° and the maximum lift coefficient reduces about 50% which is hazardous for an airplane. While horn ice leads to a stall angle drop of about 4° and a maximum lift coefficient reduction to 21%, runback ice has the least effect on the flow pattern around the airfoil and the aerodynamic coefficients so as the stall angle decreases 2° and the maximum lift reduces about 8%.

Numerical Investigation on Uniformity of Heat Distribution of Swirl Anti-Icing System

Heat distribution characteristics e.g. temperature, the local Nusselt number, local heat flux etc., contribute to ice accretion on all critical surfaces of aircraft. The non-uniformity in heat distribution at these surfaces results in low temperatures and runback ice accretion downstream from these surfaces areas. Most of the previous studies merely extracted local temperature or local Nusselt number profiles at certain locations/cross-sectional planes on the surface to describe the uniformity of the temperature, which of two does not represent the overall temperature distribution on the entire surface. This paper proposes using the Coefficient of temperature deviation (Ctem dev), which is obtained using Statistical Quality Control (SQC) and statistical methods, as the measurement of the anti-icing system’s temperature uniformity. The discussion covers the uniformity of the temperature distribution in a nacelle lip, which is produced by Swirl Anti-Icing (SAI) system. This paper also numerically investigates the effects of an SAI system's nozzle length and nozzle rotation angle on the temperature uniformity over several Reynolds numbers (Re) in a D-chamber, ranging between 7.5 x 104 and 4.1 x 105. At Re= 7.58 x 104 and velocity of free stream = 48.75 m/s, Ctem dev decreases by 13.7% as the Nozzle Direction Angle (NRA) changes from 0⁰ to 13⁰, which indicates that changing nozzle direction significantly improves the uniformity of the temperature distribution on a nacelle lip surface. However, according to the numerical results, Ctem dev is insensitive to nozzle length. The results also show that the uniformity of the temperature distribution improves as the Re increases, for NRA values of both 0⁰ and 13⁰.

Runback ice formation mechanism on hydrophilic and superhydrophobic surfaces

Experimental results aimed at understanding the different mechanisms of ice accretion on surfaces with different wettabilities are presented. Ice accretion was studied on a hydrophilic and a superhydrophobic surface of NACA 0021 airfoils, during tests inside an icing wind tunnel. Visualization of ice accretion was performed using an infrared camera, which allows an enhanced view of liquid water and ice present on the airfoil surface compared to the optical imaging. The use of the infrared camera permitted identification of different ice shapes and ice formation mechanisms on the test articles: on the hydrophilic sample, a compact ice front accreted, whereas on the superhydrophobic sample, small isolated ice islands formed. In addition, the ice on the superhydrophobic sample was more susceptible to be shed from the surface, as shown by shedding of several ice islands due to aerodynamic drag during tests. Combined analysis of infrared camera images and optical images confirmed that the fraction of the airfoil covered by ice decreases for the superhydrophobic sample, for a two minute test, for all the input heat power tested.

Numerical simulation of an airfoil electrothermal anti-icing system

The mathematical models and a numerical code for numerical simulation of a thermal anti-icing system are presented in this article. Mass conservation is applied to the runback water flow. An energy balance is imposed on the airfoil skin including the water flow. The heat transfer coefficient distributions are obtained using the boundary layer integral method. The external flowfield and the local water collection efficiency data are predicted using an Eulerian method, based on a computation fluid dynamic code and its user-defined functions. Given the input of the electrical power density distribution, the numerical code is able to calculate airfoil equilibrium surface temperature, mass flux of runback water, runback ice mass flux, and range if happens. A user interface is developed to integrate the computation fluid dynamic code to achieve a method for the analysis of a thermal anti-icing system. All the numerical results are compared with both experimental data and other numerical results presented in the literature.

G050074 軸流ファンヘの着氷に関する実験的研究 : 着氷形態の影響について

Ice accretion on rotor blades with three forms of icing such as soft rime, hard rime and glaze ice was investigated experimentally with a commercially available axial fan. In the experiments, super cooled water generated by a spray nozzle was drawn into the axial fan in a low temperature room. Results revealed that thickness of soft rime and hard rime on leading edge in the vicinity of a hub was small, and a hump was made at a distance with 10% length of a blade width from a hub. And a runback ice accretion which was suggested by radial line pattern was observed with glaze ice, and mass of the ice accretion became large in order of soft rime, hard rime and glaze ice.

Experimental Setup for the Study of Runback Ice at Full Scale

a = length in front of shape’s ridge, m B = global catch efficiency b = length of shape’s ridge, m c = chord, m D = distance between end of limit zone and ice shape, m d = height of shape’s ridge, m H = heating system configuration LWC = liquid water content, g m 3 Ma = Mach number MVD = median volumetric diameter, m Re = Reynolds number U1 = freestream velocity, m s 1 T = tunnel TS = surface temperature, C Ts = static temperature, C T1 = freestream temperature, C X = position along chord, m = angle between normal to surface and shape’s ridge