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Active flow control on the NACA 0024 airfoil defined as suction-injection jet at the chord-based Reynolds number of 1.5 × 1e + 5 is studied.

The three-dimensional incompressible unsteady Reynolds-averaged Navier–Stokes equations with the SST k-ω turbulence model are used to study the effects of coflow-jet (CFJ) on the dynamic and static stall phenomena. CFJ implementation is conducted with several momentum coefficients to investigate their turnover. Furthermore, the current work intends to analyze the CFJ performance by varying the Reynolds number and jet momentum coefficient and comparing all states to the baseline airfoil, which has not been studied in prior research investigations.

It is observed that at the momentum coefficient (Cµ) of 0.06, the lift coefficients at low attack angles (up to a = 15) dramatically increase. Furthermore, the dynamic stall at the given Reynolds number and with the lowered frequency of 0.15 is explored. In the instance of Cµ = 0.07, the lift coefficient curve does not show a noticeable stall feature compared to Cµ = 0.05, suggesting that a more powerful stronger jet can entirely control the dynamic stall.

Furthermore, the current work intends to analyze the CFJ performance by varying the jet momentum coefficient and comparing all states to the baseline airfoil, which has not been studied in prior research investigations.

This paper aims to conduct, a detailed investigation of various Reynolds averaged Navier–Stokes (RANS) models to study their performance in attached and separated flows. The turbulent flow over two airfoils, namely, National Advisory Committee for Aeronautics (NACA)-0012 and National Aeronautics and Space Administration (NASA) MS(1)-0317 with a static stall setup at a Reynolds number of 6 million, is chosen to investigate these models. The pre-stall and post-stall regions, which are in the range of angles of attack 0°–20°, are simulated.

RANS turbulence models with the Boussinesq approximation are the most commonly used cost-effective models for engineering flows. Four RANS models are considered to predict the static stall of two airfoils: Spalart–Allmaras (SA), Menter’s k – ω shear stress transport (SST), k – kL and SA-Bas Cakmakcioglu modified (BCM) transition model. All the simulations are performed on an in-house unstructured-grid compressible flow solver.

All the turbulence models considered predicted the lift and drag coefficients in good agreement with experimental data for both airfoils in the attached pre-stall region. For the NACA-0012 airfoil, all models except the SA-BCM over-predicted the stall angle by 2°, whereas SA-BCM failed to predict stall. For the NASA MS(1)-0317 airfoil, all models predicted the lift and drag coefficients accurately for attached flow. But the first three models showed even further delayed stall, whereas SA-BCM again did not predict stall.

The numerical results at high Re obtained from this work, especially that of the NASA MS(1)-0317, are new to the literature in the knowledge of the authors. This paper highlights the inability of RANS models to predict the stall phenomenon and suggests a need for improvement in modeling flow physics in near- and post-stall flows.

This paper aims to deal with the numerical investigation of laminar separation bubble (LSB) characteristics (length and height of the bubble) of SS007 airfoil at the chord Reynolds number of Re_c = 0.68 × 10⁵ to 10.28 × 10⁵.

The numerical simulations of the flow around SS007 airfoil were carried out by using the commercial fluid dynamics (CFD) software, ANalysis system (ANSYS) 15. To solve the governing equations of the flow, a cell-centred control volume space discretisation approach is used. Wind tunnel experiments were conducted at the chord-based Reynolds number of Re_c = 1.6 × 10⁵ to validate the aerodynamic characteristics over SS007 airfoil.

The numerical results revealed that the LSB characteristics of a SS007 airfoil, and the aerodynamic performances are validated with experimental results. The lift and drag coefficients for both numerical and experimental results show very good correlation at Reynolds number 1.6 × 10⁵. The lift coefficient linearly increases with the increasing angle of attack (AOA) is relatively small. The corresponding drag coefficient was found to be very small. After the formation of LSB which leads to burst to cause airfoil stall, the lift coefficient decreases and increases the drag coefficient.

Low Reynolds number and LSB characteristics concept in aerodynamics is predominant for both civilian and military applications. These include high altitude devices, wind turbines, human powered vehicles, remotely piloted vehicles, sailplanes, unmanned aerial vehicle and micro aerial vehicle. In this paper, the micro aerial vehicle flight conditions considered and investigated the LSB characteristics for different Reynolds number. To have better aerodynamic performances, it is strongly recommended to micro aerial vehicle (MAV) design engineers that the MAV is to fly at 12 m/s (cruise speed).

MAVs and unmanned aerial vehicles seem to give some of the technical challenges of nature conservation monitoring and law enforcement a versatile, reliable and inexpensive solution.

The SS007 airfoil delays the flow separation and improves the aerodynamic efficiency by increasing the lift and decreasing the drag. The maximum increase in aerodynamic efficiency is 12.5% at stall angle of attack compared to the reference airfoil at Re = 2 × 10⁵. The results are encouraging and this airfoil could have better aerodynamic performance for the development of MAV.

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.

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 describe a methodology for predicting the effects of glaze ice geometry on airfoil aerodynamic coefficients by using neural network (NN) prediction. Effects of icing on angle of attack stall are also discussed.

The typical glaze ice geometry covers ice horn leading‐edge radius, ice height, and ice horn position on airfoil surface. By using artificial NN technique, several NNs are developed to study the correlations between ice geometry parameters and airfoil aerodynamic coefficients. Effects of ice geometry on airfoil hinge moment coefficient are also obtained to predict the angle of attack stall.

NN prediction is feasible and effective to study the effects of ice geometry on airfoil performance. The ice horn location and height, which have a more evident and serious effect on airfoil performance than ice horn leading‐edge radius, are inversely proportional to the maximum lift coefficient. Ice accretions on the after‐location of the upper surface of the airfoil leading edges have the most critical effects on the airfoil performance degradation. The catastrophe of hinge moment and unstable hinge moment coefficient can be used to predict the stall effectively.

Since the simulation results of NNs are shown to have high coherence with the tunnel test data, it can be further used to predict coefficients at non‐experimental conditions.

The simulation method by using NNs here can lay the foundation of the further research about the airfoil performance in different ice cloud conditions through predicting the relations between the ice cloud conditions and ice geometry.

The complex flow behavior over an oscillating aerodynamic body, e.g. a helicopter rotor blade, a rotating wind turbine blade or the wing of a maneuvering airplane involves combinations of pitching and plunging motions. As the parameters of the problem (Re, St and phase difference between these two motions) vary, a quasi-steady analysis fails to provide realistic results for the aerodynamic response of the moving body, whereas this study aims to provide reliable experimental data.

In the present study, a pitching and plunging mechanism was designed and built in a subsonic closed-circuit wind tunnel as well as a rectangular aluminum wing of a 2:1 aspect-ratio with a NACA64-418 airfoil, used in wind turbine blades. To measure the pressure distribution along the wing chord, a number of fast responding transducers were embedded into the mid span wing surface. Simultaneous pressure measurements were conducted along the wing chord for the Reynolds number of 0.85 × 10⁶ for both steady and unsteady cases (pitching and plunging). A flow visualization technique was used to detect the flow separation line under steady conditions.

Elevated pressure fluctuations coincide with the flow separation line having been detected through surface flow visualization and flattened pressure distributions appear downstream of the flow separation line. Closed hysteresis loops of the lift coefficient versus angle of attack were measured for combined pitching and plunging motions.

The experimental data can be used for improvement of unsteady fluid mechanics problem solvers.

In the present study, a new installation was built allowing the aerodynamic study of oscillating wings performing pitching and plunging motions with prescribed frequencies and phase lags between the two motions. The experimental data can be used for improvement of computational fluid dynamics codes in case that the examined aerodynamic body is oscillating.

The purpose of this paper is to explore the effects of the camber on gliding and hovering performance of two-dimensional corrugated airfoils. While the flying mechanism of natural flyers remains a myth up to nowadays, the simulation serves as a minor step toward understanding the steady and unsteady aerodynamics of the dragonfly flight.

The lattice Boltzmann method is used to simulate the flow past the cambered corrugated dragonfly airfoil at low Reynolds numbers. For gliding flight, the maximum camber, the distance of the location of maximum camber point from the leading edge and Reynolds number are regarded as control variables; for hovering flight, the maximum camber, the flapping amplitude and trajectory are considered as control variables. Then corresponding simulations are performed to evaluate the implications of these factors.

Greater gliding ratio can be reached by increasing the maximum camber of the dragonfly wing section. When the location of the maximum camber moves backward along the wing chord, large scale flow separation can be delayed. These two effects result in better gliding performances. For hovering performances, it is found that for different flapping amplitudes along an inclined plane, the horizontal force exerted on the airfoils increases with the camber, and the drag growths first but then drops. It is also found that the elliptic flapping trajectory is most sensitive to the camber of the cambered corrugated dragonfly wing section.

The effects of the camber on gliding and hovering performance of the cambered dragonfly wing section are explored in detail. The data obtained can be helpful when designing micro aerial vehicles.

The purpose of the paper is to analyse different post-stall models, originally developed for use in wind turbine codes, and extend their use to the propeller performance prediction.

Different post-stall methods available in the literature were implemented in JBLADE software. JBLADE contains an improved version of Blade Element Momentum theory, and it is appropriate for the design and analysis of different propellers in off-design conditions.

The preliminary analysis of the results shows that the propeller performance prediction can be improved using these implemented post-stall models. However, there is a lack of accuracy in the performance prediction of some propellers. Further comparisons including distribution of forces along the blade may help to better understand this inaccuracy of the models, and it will be studied in future work.

The work has extended the use of the post-stall models to the propeller performance prediction codes. It is shown that these models can be used to obtain an improved prediction of the propeller’s performance.

Rainfall is one of the main atmospheric conditions that significantly affect the aerodynamic performance of the low Reynolds number flights. In this paper, the adverse effects of rain on the aerodynamic performance of a two-dimensional (2D) airfoil with a chord-based low Reynolds number of 2 × 10⁵ and the mini-unmanned aerial vehicle (UAV) for various flight conditions, i.e. 0°–40° at Mach number 0.04 were studied numerically. The purpose of this study is to explore the aerodynamic penalties that affect the liquid water content (LWC = 5.33) of the airfoil and UAV performance in rain under different flying conditions.

The Eulerian–Lagrangian two-phase flow method is adopted to simulate the rain environment over an airfoil and mini-UAV aerodynamic performances. The Reynolds Averaged Navier–Stokes equations are considered to solve the time-averaged equations of motion for fluid flow.

The effect of rainfall on the airfoil and mini-UAV is studied numerically and validated experimentally. For 2D airfoil, the lift and drag coefficients for both numerical and experimental results show a very good correlation at Reynolds number 2 × 10⁵. For three-dimensional (3D) mini-UAV, the lift and drag coefficients for both numerical and experimental results show a very good correlation at Mach number 0.04. The raindrops distribution around the airfoil, premature trailing edge separation, boundary-layer velocity profiles at five different chord positions (i.e. LE, 0.25c, 0.5c, 0.75c and 0.98c) on the upper surface of the airfoil, water film height and the location of rivulet formation on the upper surface of the airfoil are also presented.

For 2D airfoil, the recorded maximum variation of the coefficient of lift and lift-to-drag (L/D) ratio is observed to be 5.33% at an 8° and 10.53% at a 4° angle of attack (AOA) between numerical and experimental results under the influence of rainfall effect for LWC = 5.33. The L/D ratio percentage degradation is seen to be 61.9% at an AOA of 0°–2° for the rain environment. For 3D mini-UAV, the recorded maximum variation of the coefficient of lift and L/D ratio are observed to be 2.84% and 4.60% at a 30° stall AOA under the influence of rainfall effect for LWC = 5.33. The numerical results are impressively in agreement with the experimental results. UAV designers will benefit from the findings presented in this paper. This will be also helpful for training the pilots to control the airplanes in a rain environment.