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This paper aims to achieve an optimum flow separation control over the airfoil using a passive flow control method by introducing a bio-inspired nose near the leading edge of the National Advisory Committee for Aeronautics (NACA) 4 and 6 series airfoil. In addition, to find the optimised leading edge nose design for NACA 4 and 6 series airfoils for flow separation control.

Different bio-inspired noses that are inspired by the cetacean species have been analysed for different NACA 4 and 6 series airfoils. Bio-inspired nose with different nose length, nose depth and nose circle diameter have been analysed on airfoils with different thicknesses, camber and camber locations to understand the aerodynamic flow properties such as vortex formation, flow separation, aerodynamic efficiency and moment.

The porpoise nose design that has a leading edge with depth = 2.25% of chord, length = 0.75% of chord and nose diameter = 2% of chord, delays the flow separation and improves the aerodynamic efficiency. Average increments of 5.5% to 6° in the lift values and decrements in parasitic drag (without affecting the pitching moment) for all the NACA 4 and 6 series airfoils were observed irrespective of airfoil geometry such as different thicknesses, camber and camber location.

The two-dimensional computational analysis is done for different NACA 4 and 6 series airfoils at low subsonic speed.

This design improves aerodynamic performance and increases the structural strength of the aircraft wing compared to other conventional high lift devices and flow control devices. This universal leading edge flow control device can be adapted to aircraft wings incorporated with any NACA 4 and 6 series airfoil.

The results would be of significant interest in the fields of aircraft design and wind turbine design, lowering the cost of energy and air travel for social benefits.

Different bio-inspired nose designs that are inspired by the cetacean species have been analysed for NACA 4 and 6 series airfoils and universal optimum nose design (porpoise airfoil) is found for NACA 4 and 6 series airfoils.

The purpose of this paper is to achieve an optimum flow separation control over the airfoil using passive flow control method by introducing bio-inspired nose near the leading edge of the NACA 2412 airfoil.

Two distinguished methods have been implemented on the leading edge of the airfoil: forward facing step, which induces multiple accelerations at low angle of attack, and cavity/backward facing step, which creates recirculating region (axial vortices) at high angle of attack.

The porpoise airfoil (optimum bio-inspired nose airfoil) delays the flow separation and improves the aerodynamic efficiency by increasing the lift and decreasing the parasitic drag. The maximum increase in aerodynamic efficiency is 22.4 per cent, with an average increase of 8.6 per cent at all angles of attack.

The computational analysis has been done for NACA 2412 airfoil at low subsonic speed.

This design improves the aerodynamic performance and increases structural strength of the aircraft wing compared to other conventional high-lift devices and flow-control devices.

Different bio-inspired nose designs which are inspired by the cetacean species have been analysed for NACA 2412 airfoil, and optimum nose design (porpoise airfoil) has been found.

The purpose of this paper is to investigate the effect of surface protrusions on the flow unsteadiness of NACA 0012 at a Reynolds number of 100,000.

Effect of protrusions is investigated through numerical simulation of two-dimensional Navier–Stokes equations using a finite volume solver. Turbulent stresses are resolved through the transition Shear stress transport (four-equation) turbulence model.

The small protrusion located at 0.05c and 0.1c significantly improve the lift coefficient by up to 36% in the post-stall regime. It also alleviates the leading edge stall. The larger protrusions increase the drag significantly along with significant degradation of lift characteristics in the pre-stall regime as well. The smaller protrusions also increase the frequency of the vortex shedding.

The effect of macroscopic protrusions or deposits in rarely investigated. The delay in stall shown by smaller protrusions can be beneficial to micro aerial vehicles. The smaller protrusions increase the frequency of the vortex shedding, and hence, can be used as a tool to enhance energy production for energy harvesters based on vortex-induced vibrations and oscillating wing philosophy.

This paper aims to model the aerodynamic flow characteristics of NACA0010 for various angle of attacks including stall for incompressible flows using panel methods. This paper also aims to quantify the surface pressure distribution on streamlined bodies and validate the results with analytical Jukouwski method and inverse panel methods that can predict the aerodynamic flow behaviour using the geometric iteration approach.

The 2 D panel method was implemented in Qblade software v.06 which uses the fundamental panel method which rely on source strengths and influence coefficients to determine the velocity and pressure fields on the surface. The software implements the boundary layer or viscous effects to determine the influence on aerodynamic performance at various angles of attack. Jukouwski method is also evaluated for predicting aerodynamic characteristics and is based on the geometric iteration approach. Then complex aerodynamic flow potentials are determined based on the source strengths which are used to predict the pressure and velocity fields.

At low to moderate angles of attack, panel and Jukouwski methods predict similar results for surface pressure coefficients comparable to Hess and Smith inverse method. In comparison to panel method, results from the Jukouwski mapping method predicted the pressure coefficient conservatively for the same free stream conditions. With increase in Reynolds number, lift coefficient and aerodynamic performance improved significantly for un-tripped aerofoil when stall angle is approached when compared to tripped aerofoil.

This study demonstrated that panel methods have higher efficacy in terms of computational time or resources and thus can provide benefits to many real-world aircraft or aerospace design applications.

Even though panel and Jukouwski methods have been studied extensively in the past, this paper demonstrates the efficacy of both methods for modelling aerodynamic flows that range between moderate to high Reynolds number which are critical for many aircraft applications. Both methods have been validated with analytical and inverse design methods which are able to predict aerodynamic flow characteristics for simple bluff bodies, streamlined aerofoils as well as bio-inspired corrugated aerofoils.

The aims of this study were to investigate the effect of using a wall‐normal, 2D micro zero‐net‐mass‐flux (ZNMF) jet located at the leading edge of a NACA 0015 airfoil to actively control flow separation and enhance lift.

Experiments were conducted over a two‐dimensional airfoil in a water tunnel at a Reynolds number of 3.08 × 10⁴ for the parametric investigation and the detailed multigrid cross‐correlation digital particle image velocimetry (MCCDPIV) measurements. Flow visualisation experiments were carried out at a lower Reynolds number of 1.54 × 10⁴.

The largest lift increase was observed when a non‐dimensional frequency of 1.3 and an oscillatory momentum blowing coefficient of 0.14 per cent was employed. Under these forcing conditions the stall angle of the airfoil was mitigated from an angle of attack of 10^o to one of 18^o, resulting in a maximum lift coefficient increase of 46 per cent above the uncontrolled lift coefficient. Planar laser induced fluoroscopy and MCCDPIV revealed that the lift increments were the result of the generation of a train of large‐scale, spanwise lifting vortices that convected over the suction surface of the airfoil. The presence of these structures resulted in the flow seemingly remaining attached to the upper surface of the airfoil for a wider range of angles of attack.

This study is significant as it provides quantitative experimental data, which clearly demonstrates the effectiveness of a 2D micro ZNMF jet in controlling flow separation of a NACA 0015 airfoil at high angles of attack and thus, enhancing lift. Furthermore, the flow visualisations and MCCDPIV measurements have provided insight into the mechanisms responsible for the improvement in lift. This new understanding has applications beyond the NACA 0015 airfoil used in this study.

The purpose of this paper is to use dynamic meshing to perform CFD analyses of a NACA 0012 airfoil fitted with a morphing trailing edge (TE) flap when it undergoes static and time-dependent morphing. The steady CFD predictions of the original and morphing airfoils are validated against published data. The study also investigates an airfoil with a hinged TE flap for aerodynamic performance comparison. The study further extends to an unsteady CFD analysis of a dynamically morphing TE flap for proof-of-concept and also to realise its potential for future applications.

An existing parametrization method was modified and implemented in a user-defined function (UDF) to perform dynamic meshing which is essential for morphing airfoil unsteady simulations. The results from the deformed mesh were verified to ensure the validity of the adopted mesh deformation method. ANSYS Fluent software was used to perform steady and unsteady analysis and the results were compared with computational predictions.

Steady computational results are in good agreement with those from OpenFOAM for a non-morphing airfoil and for a morphed airfoil with a maximum TE deflection equal to 5 per cent of the chord. The results obtained by ANSYS Fluent show that an average of 6.5 per cent increase in lift-to-drag ratio is achieved, compared with a hinged flap airfoil with the same TE deflection. By using dynamic meshing, unsteady transient simulations reveal that the local flow field is influenced by the morphing motion.

An airfoil parametrisation method was modified to introduce time-dependent morphing and used to drive dynamic meshing through an in-house-developed UDF. The morphed airfoil’s superior aerodynamic performance was demonstrated in comparison with traditional hinged TE flap. A methodology was developed to perform unsteady transient analysis of a morphing airfoil at high angles of attack beyond stall and to compare with published data. Unsteady predictions have shown signs of rich flow features, paving the way for further research into the effects of a dynamic flap on the flow physics.

There is an increasing interest in airfoils that modify their shape to adapt at the flow conditions. As an example of application, the authors search the optimal 4-digit NACA airfoil that maximizes the lift-over-drag ratio for a constant lift coefficient of 0.6, from Re = 10⁴ to 3 × 10⁶.

The authors consider a γ−Re_θt transition model and a κ – ω SST turbulence model with a covariance matrix adaptation evolutionary optimization algorithm. The shape is adapted by radial basis functions mesh morphing using four parameters (angle of attack, thickness, camber and maximum camber position). The objective of the optimization is to find the airfoil that enables a maximum lift-over-drag ratio for a target lift coefficient of 0.6.

The computation of the optimal airfoils confirmed the expected increase with Re of the lift-over-drag ratio. However, although the observation of efficient biological fliers suggests that the thickness increases monotonically with Re, the authors find that it is constant but for a 1.5 per cent step increase at Re = 3 × 10⁵.

The authors propose and validate an efficient high-fidelity method for the shape optimization of airfoils that can be adopted to define robust and reliable industrial design procedures.

The authors show that the difference in the numerical error between two-dimensional and three-dimensional simulations is negligible, and that the numerical uncertainty of the two-dimensional simulations is sufficiently small to confidently predict the aerodynamic forces across the investigated range of Re.

This paper aims to study the extended trailing edge airfoil for a range of angle of attack at different intensities of turbulence.

In this paper, an experimental study on NACA 0020 airfoil with thin extended trailing edge modification of amplitude of h = 0.1c, 0.2c and 0.3c at the Reynolds number of 2.14 × 10⁵ are tested. The research was carried out for an angle of attack ranging from 0° = α = 35° for the turbulence intensity of 0.3%, 3%, 5%, 7% and 12%. From the experimental readings, the surface pressures are scanned using a Scanivalve (MPS2464) pressure scanner for a sampling frequency of 700 Hz. The scanned pressures are converted to aerodynamic force coefficient and the results are combined and discussed.

The airfoil with the extended trailing edge will convert the adverse pressure gradient to a plateau pressure zone, indicating the delayed flow separation. The CL value at higher turbulence intensity (TI = 12%) for the extended trailing edge over perform the base airfoil at the post-stall region. The maintenance of flow stability is observed from the spectral graph.

A thin elongated trailing edge attached to the conventional airfoil serves as a flow control device by delaying the stall and improving the lift characteristics. Additionally, extending the airfoil's trailing edge helps to manage the performance of the airfoil even at a high level of turbulence.

Distinct from existing studies, the presented results reveals how the extended trailing edge attached to the airfoil performs in the turbulence zone ranging from 0.3% to 12% of TI. The displayed pressure distribution explains the need for increasing trailing edge amplitude (h) and its impact on flow behaviour. The observation is that extended trailing edge airfoil bears to maintain the performance even at higher turbulence region.

In this paper, the applicability of shear stress transport k-ω model along with the intermittency concept has been investigated over pitching airfoils to capture the laminar separation bubble (LSB) position and the boundary layer transition movement. The effect of reduced frequency of oscillations on boundary layer response is also examined.

A two-dimensional computational fluid dynamic code was developed to compute the effects of unsteadiness on LSB formation, transition point movement, pressure distribution and lift force over an oscillating airfoil using transport equation of intermittency accompanied by the k-ω model.

The results indicate that increasing the angle of attack over the stationary airfoil causes the LSB size to shorten, leading to a rise in wall shear stress and pressure suction peak. In unsteady cases, both three- and four-equation models are capable of capturing the experimentally measured transition point well. The transition is delayed for an unsteady boundary layer in comparison with that for a static airfoil at the same angle of attack. Increasing the unsteadiness of flow, i.e. reduced frequency, moves the transition point toward the trailing edge of the airfoil. This increment also results in lower static pressure suction peak and hence lower lift produced by the airfoil. It was also found that the fully turbulent k-ω shear–stress transport (SST) model cannot capture the so-called figure-of-eight region in lift coefficient and the employment of intermittency transport equation is essential.

Boundary layer transition and unsteady flow characteristics owing to airfoil motion are both important for many engineering applications including micro air vehicles as well as helicopter blade, wind turbine and aircraft maneuvers. In this paper, the accuracy of transition modeling based on intermittency transport concept and the response of boundary layer to unsteadiness are investigated.

As a conclusion, the contribution of this paper is to assess the ability of intermittency transport models to predict LSB and transition point movements, static pressure distribution and aerodynamic lift variations and boundary layer flow pattern over dynamic pitching airfoils with regard to oscillation frequency effects for engineering problems.

One of the best methods to improve wind turbine aerodynamic performance is modification of the blade’s airfoil. The purpose of this paper is to investigate the effects of gurney flap geometry and its oscillation parameters on the pitching NACA0012 airfoil.

This numerical solution has been carried out for different cases of gurney flap mounting angles, heights, reduced frequencies and oscillation amplitudes, then the results were compared to each other. The finite volume method was used for the discretization of the governing equations, and the PISO algorithm was used to solve the equations. Also, the “SST” was adopted as the turbulence model in the simulation.

In this paper, the different parameters of gurney flap were investigated. The results showed that the best range of gurney flap height are between 1 and 3.2% of chord and the best ratio of lifting to drag coefficient is achieved in gurney flap with an angle of 90° relative to the chord direction. The dynamic stall angle of the airfoil with gurney flap decreases were compared to without gurney flap. Earlier LEV formation can be one of the main reasons for decreasing the dynamic stall angle of the airfoil with gurney flap. Increasing the reduced frequency and oscillation amplitude causes rising of maximum lift coefficient and consequently lift curve slope. Moreover, gurney flap with mounting angle has a lower hinge moment than the gurney flap without mounting angle but with the same vertical axis length. So, there is more complexity in structural design concerning the gurney flap without mounting angle.

Improving aerodynamic efficiency of airfoils is vital for obtaining more output power in VAWTs. Gurney flaps are one of the best mechanisms to increase the aerodynamic performance of the airfoil and increases the efficiency of VAWTs.

Investigating the hinge moment on the connection point of the airfoil, gurney flap and try to compare the gurney flap with and without angle.