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Fluidic oscillators are used to overcome fluid flow separation on the upper side of the airfoil. The AoA chosen is for stall conditions, namely 16°, 18° and 20°. The type of fluidic oscillator used is a two-feedback channel fluidic oscillator. The chosen element is a triangle because it is easy to implement in complex geometries such as fluidic oscillators. The algorithm used is PISO. The governing equation for solving problems is URANS. URANS is then combined with the k-omega SST turbulence model. A low Reynolds number of 48000 was chosen to simulate the NACA 0015 airfoil. This Reynolds number is calculated based on the chord length of the airfoil. Variations were also made to the fluidic oscillator velocity-inlet value. The velocity-inlet variations given are 25 m/s, 35 m/s, and 45 m/s. The increase in Cl will be more significant if the selected inlet velocity is higher. At an inlet velocity of 45 m/s, the average increase produced was 38.53%. Double FO experienced an increase of 21.99%. Drag reduction is another parameter used to assess the aerodynamic performance of an airfoil. The average drag reduction for a velocity-inlet of 45 m/s is 12.54%. Double FO produces an average drag reduction of 6.80%.

The objective of this study is to develop a genetic algorithm that uses the IGP parameterization to increase the lift coefficient (CL) of three airfoils to be used on wings of unmanned aerial vehicles (UAVs). The geometry of three baseline airfoils was modified by developing a genetic algorithm that operates with the IGP parameterization and performs the aerodynamic analysis using XFOIL in the MATLAB environment. Subsequently, a numerical model was made for each baseline and optimized airfoil using a commercial computational fluid dynamics (CFD) code to analyze the behavior of the lift coefficient. An increase in the average CL was obtained for the Eppler 68, MH 70, and Wortmann FX 60-126 airfoils for angles of attack ranging from 0 to 10, obtaining increments of 17.243%, 14.967%, and 10.708%, respectively. Additionally, an average 5.027% uncertainty was obtained in lift coefficient calculations between XFOIL and CFD. The utility of the IGP method and genetic algorithms for parameterizing and optimizing airfoils was demonstrated. In addition, airfoils could be tailored for a specific UAV depending on the mission profile. Volume 2 of this study will include experimental data from wind tunnel.

Effect of airfoil thickness on onset of dynamic stall is investigated using large eddy simulations at chord-based Reynolds number of 200 000. Four symmetric NACA airfoils of thickness-to-chord ratios of 9 %, 12 %, 15 % and 18 % are studied. The three-dimensional Navier–Stokes solver, FDL3DI is used with a sixth-order compact finite difference scheme for spatial discretization, second-order implicit time integration and discriminating filters to remove unresolved wavenumbers. A constant-rate pitch-up manoeuver is studied with the pitching axis located at the airfoil quarter chord. Simulations are performed in two steps. In the first step, the airfoil is kept static at a prescribed angle of attack ( $=4^{\\circ }$ ). In the second step, a ramp function is used to smoothly increase the pitch rate from zero to the selected value and then the pitch rate is held constant until the angle of attack goes past the lift-stall point. The solver is verified against experiments for flow over a static NACA 0012 airfoil. Static simulation results of all airfoil geometries are also compared against XFOIL predictions with a generally favourable agreement. FDL3DI predicts two-stage transition for thin airfoils (9 % and 12 %), which is not observed in the XFOIL results. The dynamic simulations show that the onset of dynamic stall is marked by the bursting of the laminar separation bubble (LSB) in all the cases. However, for the thickest airfoil tested, the reverse flow region spreads over most of the airfoil and reaches the LSB location immediately before the LSB bursts and dynamic stall begins, suggesting that the stall could be triggered by the separated turbulent boundary layer. The results suggest that the boundary between different classifications of dynamic stall, particularly leading edge stall versus trailing edge stall, is blurred. The dynamic-stall onset mechanism changes gradually from one to the other with a gradual change in some parameters, in this case, airfoil thickness.

NACA 0012 symmetric airfoil is commonly used for the aerodynamic analysis of aircraft. This research work investigates modifications in the airfoil design to improve the lift-drag ratios. The initial phase involves computational analysis on the two-dimensional airfoil and later, the obtained results were validated by comparing them to standard data obtained by calculations. The modified airfoil designs are stepped trailing edge, tubercle leading edge and combination of both. The design alterations implemented was based on insights gained from various simulations with the adjustment in angle of attack ranging from 0° to 8°. The stepped design in trailing edge generated less drag and produced less lift. Whereas, the tubercle in leading edge of airfoil exhibited more drag and lift forces. Although generation of lift is favourable for an aircraft, it is of paramount importance to reduce its drag counterpart for efficient flying. Hence, the characteristics of tubercle leading edge and stepped trailing edge airfoil was incorporated into a single airfoil and their aerodynamic performance results were explained in this paper.

An objective of mechanical design is to obtain a shape that satisfies specific requirements. In the present work, we achieve this goal using a conditional variational autoencoder (CVAE). The method enables us to analyze the relationship between aerodynamic performance and the shape of aerodynamic parts, and to explore new designs for the parts. In the CVAE model, a shape is fed as an input and the corresponding aerodynamic performance index is fed as a continuous label. Then, shapes are generated by specifying the continuous label and latent vector. When CVAE is applied to mechanical design, it is desired to draw shapes that reproduce the specified aerodynamic performance. In ordinal CVAE, the model is trained to minimize reconstruction loss and latent loss, and it is usually optimized considering the sum of these losses. However, the present study shows that the optimal network is not always optimal in terms of reproducing the aerodynamic performance. The proposed method is verified using two numerical examples: a two-dimensional (2D) airfoil and a turbine blade. In the airfoil example, we demonstrate the effects of latent dimension, and in the turbine design example, we demonstrate that the proposed method can be applied to a real turbine design problem and reduce the design time.

Based on the NACA0012 airfoil, different types of protrusions namely serrated protrusions and arc-shaped protrusions are set near the trailing edge. Both new airfoils have a certain decrease in flow performance however there is a significant reduction in aerodynamic noise of airfoils. This is because the protrusions damage the vortex near the trailing edge. In addition, the noise reduction of serrated protrusions is more significant than t of arc-shaped protrusions, about 2.7dB.

An aircraft in practice serves under extreme flight conditions that will have a substantial impact on its flight safety. Understanding dynamics of airfoil structure of an aircraft subjected to severe load conditions is thus extremely valuable and necessary. In this study, we will explore the complicated dynamical behaviors of a conceptual airfoil excited by an external harmonic force and an extreme random load. Importantly, such an extreme random load is portrayed by a non-Gaussian Lévy noise with a heavy-tailed feature. Bistable behaviors of the deterministic airfoil system are performed firstly from amplitude–frequency response and basin of attraction. Then, the effects of the extreme random load on the airfoil system are thoroughly investigated. Interestingly, within the bistable regime, the extreme random load can lead to stochastic transition and stochastic resonance. Due to its heavy-tailed nature, the Lévy noise would increase the possibility of a highly unexpected stochastic transition behavior between desirable low-amplitude and catastrophic high-amplitude oscillations compared with the Gaussian scenario. Such vibration patterns might damage or destroy the airfoil structure, which will put an aircraft in great danger. All the findings would be helpful in ensuring the flight safety and enhancing the strength and reliability of airfoil structure operating at extreme flight conditions.

The NACA4412 airfoil is often chosen as the main airfoil profile for ground effect airfoils, which has a large leading-edge radius, large curvature, and flat lower airfoil surface, and has good aerodynamic performance. In this paper, a special “anti-S” airfoil profile is proposed, and an anti-S airfoil, NACAM27, is selected to be compared with the NACA4412 airfoil, which is commonly used in the field of ground effect airfoils, to study the aerodynamic performances of the two airfoils by numerical simulation under different ground effect conditions, and to optimize the better airfoil to obtain the optimal airfoil. Finally, the geostatic stability of the airfoil is analyzed. The results discuss the aerodynamic characteristics of the anti-S airfoil and the NACA4412 under subsonic conditions, the anti-S airfoil has a higher lift-to-drag ratio, and the static stability margin is easier to predict.

The dynamic performance of an oscillating airfoil subjected to the wake of a circular cylinder is studied in this paper. Two-dimensional numerical simulations are conducted at Re = 1100 by using an immersed boundary method together with the adaptive mesh refinement technique. The effects of two parameters, the gap between the cylinder and the airfoil and the oscillation frequency, are of particular interest to the present study. Therefore, dynamic responses are presented as functions of the two parameters, including the fluid forces, the associated frequency characteristics, and the energy exchange between the airfoil and the fluid. The results show that the cylinder wake can significantly reduce the drag as well as the energy extraction of the lift on the airfoil. Different synchronization behaviors between the airfoil’s oscillation and the wake pattern have been observed for some specific cases, i.e., the 1:1, 1:2 and 1:3 patterns. Remarkably, the 1:2 pattern is associated with an asymmetric vortex shedding pattern, which can further result in non-zero time-averaged lift and moment on the airfoil even though both the upstream vortices from the cylinder and the oscillation of the airfoil are periodic. Due to the strong nonlinear interaction between the cylinder wake and the airfoil’s oscillation, new frequency branches associated with nonlinear frequency superposition are formed in the responses of the airfoil and their characteristics have been demonstrated. The present study also finds that the oscillation amplitudes are important in determining the synchronization behavior.

A low-order physics-based model to simulate the unsteady flow response to airfoils undergoing large-amplitude variations of the camber is presented in this paper. Potential-flow theory adapted for unsteady airfoils and numerical methods using discrete-vortex elements are combined to obtain rapid predictions of flow behavior and force evolution. To elude the inherent restriction of thin-airfoil theory to small flow disturbances, a time-varying chord line is proposed in this work over which to satisfy the appropriate boundary condition, enabling large deformations of the camber line to be modeled. Computational fluid dynamics simulations are performed to assess the accuracy of the low-order model for a wide range of dynamic trailing-edge flap deflections. By allowing the chord line to rotate with trailing-edge deflections, aerodynamic loads predictions are greatly enhanced as compared to the classical approach where the chord line is fixed. This is especially evident for large-amplitude deformations.