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1098 ZHOU ET AL. (DBD) plasma generation for aircraft icing mitigation. DBD plasma actuators, which are fully electronic devices without any moving parts, have been widely used in recent years for active flow control to suppress flow separation and airfoil stalls [16–20]. A DBD plasma actuator usually features two electrodes that are attached asymmetrically on the opposite side of a dielectric barrier sheet. When a high alternating current (ac) voltage is applied to the electrodes, the air over the encapsulated electrode will be ionized and generate a streak of plasma flow [21]. In the presence of a high- intensity electric field, the ionized air will lead to a body force that acts on the surrounding air [17]. During this process, because the ambient air over the encapsulated electrode will also be heated up by the plasma [22], the thermal effect induced by DBD plasma generation can be leveraged for anti-/deicing applications. Van den Broecke [23] was the first to conduct a feasibility study to explore the effectiveness of using DBD plasma to remove ice accretion from a stationary flat plate. More recently, Meng et al. [24] performed an experimental study to use a DBD plasma actuator to remove ice accretion over a circular cylinder. Although the feasibility of using DBD plasma for anti-/deicing applications has been demonstrated in those preliminary studies with simplified test models, the effectiveness of using DBD plasma actuators embedded over an airfoil/wing surface for aircraft icing mitigation under typical glaze-/ rime-icing conditions has never been explored. In the present study, an explorative study was performed to evaluate the effectiveness of using the thermal effect induced by DBD plasma generation for aircraft icing mitigation. The experimental study was performed in an icing research tunnel available at Iowa State University (i.e., ISU-IRT). A NACA 0012 airfoil/wing model embedded with DBD plasma actuators was installed inside the test section of the ISU-IRT under typical glaze- and rime-icing conditions pertinent to aircraft inflight icing phenomena. During the experiments, while a high-speed imaging system was used to record the dynamic ice accretion and water runback process over the surface of the airfoil/wing model, with and without switching on the DBD plasma actuators, an infrared thermal imaging system was used to map the corresponding surface temperature distributions over the airfoil surface simultaneously. The effectiveness of using the thermal effect induced by plasma generation for aircraft icing mitigation was examined in detail based on the side-by-side comparisons of the measurement results for the plasma-on case against those of the plasma-off case under the same icing conditions. The quantitative surface temperature measurement results were correlated with the acquired ice-accretion images to elucidate the underlying physics. II. Experimental Setup and Test Model As shown schematically in Fig. 1, the experimental study was performed in the icing research tunnel available at the Aerospace Engineering Department of Iowa State University. The icing research tunnel has a test section of 2.0 m in length by 0.4 m in width by 0.4 m in height, with four optically transparent sidewalls. The ISU-IRT has a capacity of generating a maximum wind speed of 60 m∕s and an airflow temperature of −25°C. An array of pneumatic atomizer/spray nozzles are installed at the entrance of the contraction section upstream of the test section to inject microsized water droplets (10 ∼ 100 μm in size) into the airflow. The median volume diameter of the droplets is approximately 40 μm in the present study. By manipulating the water flow rate (Q) through the water spray nozzles, the liquid water content level of the incoming airflow in the test section of the ISU-IRT can be adjusted to a desirable level, where ρ is the water density, Q is the water flow rate, A is the test section area, and U∞ is the airflow incoming speed. In summary, the ISU-IRT can be used to simulate the aircraft inflight icing phenomena over a range of icing conditions (i.e., from dry rime- to extremely wet glaze-ice conditions). In the present study, although the freestream velocity U∞ of the incoming airflow is kept constant at U∞ 40 m∕s, the LWC level and temperature T∞ of the incoming airflow are varied from 1.0 to 3.0 g∕m3 and from −5 to −15°C to simulate different (i.e., either rime or glaze) icing conditions. The airfoil/wing model used in the present study has the profile of a NACA 0012 airfoil in the cross-section with a chord length of C 150 mm and a spanwise length of L 400 mm (i.e., same dimension as the width of the ISU-IRT test section). The test model, which was manufactured by using a rapid prototyping machine (i.e., three-dimensional printing) that built the model layer by layer with a resolution of about 25 μm, is made of a hard plastic material. The airfoil/wing model was finished with a coating of primer and wet sanded to a smooth finish using 1000-grit sand paper. Supported by a stainless-steel rod, the airfoil/wing model was mounted at its quarter- chord and oriented horizontally across the middle of the test section. The angle of attack α of the airfoil/wing model was set at α −5.0 deg during the ice-accretion experiment. As shown schematically in Fig. 2, two sets of DBD plasma actuators were embedded over the pressure side of the airfoil surface. The DBD plasma actuators were arranged symmetrically to the middle span of the airfoil/wing model. During the experiment, while the DBD plasma actuators on the left side of the airfoil/wing model were turned on, the DBD plasma actuators over the right side would be kept off. The ice-accretion process over the airfoil surface for the plasma-on side (i.e., left side) would be compared side by side against that over the plasma-off side (i.e., right side) in order to evaluate the effectiveness of using the thermal effect induced by DBD plasma generation for aircraft icing mitigation under the identical icing conditions. For the present study, the DBD plasma actuators embedded over the airfoil surface consisted of copper electrodes with a thickness of about 70 μm. Three layers of Kapton film (130 μm for each layer) were used as the dielectric barrier to separate the encapsulated electrodes from the exposed electrodes. Ranging from the airfoil leading edge to around 30% of the airfoil chord, four encapsulated electrodes were distributed evenly along the airfoil chord with a separation distance of 3.0 mm. Although the buried electrodes had a chordwise width of 10.0 mm (except the first electrode at 5.0 mm in width), the exposed electrodes, which were placed above the dielectric barrier with zero overlaps against the covered electrodes, were 96 mm in length and 3.0 mm in width. Finally, the airfoil/wing model was finished with two layers of white Fig. 1 Experimental setup used in the present study. Downloaded by IOWA STATE UNIVERSITY on October 5, 2018 | http://arc.aiaa.org | DOI: 10.2514/1.J056358PDF Image | Thermal Effect by Plasma Generation to prevent Aircraft Icing
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