Electroaerodynamic (EAD) devices, which produce a propulsive force in air by electrostatic acceleration, have been demonstrated as a method of propulsion for airplanes. However, achieving sufficient thrust-to-power is a significant challenge in developing EAD aircraft which are practical. Theory predicts that devices with larger inter-electrode gap spacing will enable higher thrust-to-power, but most experimental work has been limited to gap spacings of less than 80 mm. Those studies which have investigated spacings of greater than 100 mm have found results deviating from theory, with lower thrust-to-power than predicted. We performed experiments between 50 and 300 mm gap spacing and conclude that three effects explain the discrepancy: ‘leakage current’ from the electrodes to the surroundings, which does not produce thrust but increases measured electrical power; reverse corona emission from the collecting electrode, which reduces thrust and increases power; and the electric potential of the thruster relative to its surroundings, which affects both leakage current and reverse corona emission. Our results show that if these effects are accounted for, the existing EAD theory is correct without modification beyond its previous range of validity and is applicable to wire-to-cylinder EAD devices up to at least 300 mm gap spacing. We support our experimental results with two-dimensional numerical simulations, which show that the experimental current and thrust, including effects of leakage current, can be reproduced by computation with 12% error—an important step towards numerical design and optimization. By experimentally replicating equilibrium in-flight conditions, we measure thrust-to-power in the laboratory of up to 15 N kW−1 for large gap spacing thrusters at practically useful thrust levels. This is two to three times higher than current implementations with smaller gap spacings, suggesting that large gap spacing thrusters will be suitable for future EAD-propelled flight applications at thrust-to-power competitive with or exceeding conventional propulsion.

Surface dielectric barrier discharges (SDBDs) are a type of asymmetric dielectric barrier discharge (DBD) that can be used to generate ions and produce aerodynamic forces in air. They have shown promise in a range of aerospace applications, including as actuators for solid-state aircraft control or aerodynamic enhancement and as ion sources for electroaerodynamic aircraft propulsion. However, their power draw characteristics are not well understood. Whereas existing approaches use empirical functional fits to estimate the power of specific SDBD configurations, we develop here a physics-based model for SDBD power consumption that accounts for material and geometric variation between SDBDs. The model is based on models for parallel-plate or "volume" DBDs but accounts for the "virtual electrode" resulting from changing plasma length that is particular to SDBDs. We experimentally measure the power of SDBDs of three materials, eleven thicknesses, and 29 electrical operating points to find a correlation with r 2 ¼ 0:99 (n ¼ 106) between model and experiment. We also use SDBD power measurements from four experiments in the literature and find a correlation with r 2 ¼ 0:99 (n ¼ 101) between our model and these experiments. Since we do not use any measured parameters from those experiments in our model, this suggests that our model has the ability to robustly predict the power for different SDBD construction methods and experimental techniques. Therefore, this work provides a robust method for the quantitative design and power optimization of SDBDs for a range of engineering applications, including aerospace propulsion.

Solid-state aerodynamic devices, which use electroaerodynamics (EAD) to produce a propulsive force, have the potential to make drones and airplanes significantly quieter and may provide benefits in sustainability and manufacturability. In these devices, ions are accelerated between two electrodes by an electric field, colliding with neutral air molecules and producing an ionic wind and a thrust force. The authors' previous work showed that a "decoupled" device architecture, which separates the ionization and ion acceleration processes, can increase thrust density and thrust-to-power compared to the prevailing corona-discharge-based EAD architecture, which uses a single DC potential for both processes. However, the discharge characteristics of this decoupled architecture have not been previously determined. Here, we experimentally characterize a decoupled EAD thruster with a wire-to-wire dielectric barrier discharge (DBD) ion source: an AC voltage drives the DBD, which ionizes neutral air molecules at the emitting electrode, while a separate DC voltage accelerates ions toward the collecting electrode. We determine the discharge characteristics (i.e., the DC-current-to-DC-voltage relationship) of this decoupled thruster as well as a model for the interaction between the ionization and acceleration stages: we find that the former takes the same functional form as the analytical solution for space-charge limited current in a thin collisional ion channel, whereas the latter is determined primarily by the power draw of the DBD ionization stage. We present a complete model for the thrust and power draw of decoupled EAD thrusters, enabling their quantitative design and optimization for use in aircraft propulsion and other applications.