Among semi-empirical Dynamic stall models, the boeing-vertol model uses relatively few dependent parameters determined from experimental data. Despite its simple formulation and appropriate performance, the model does not precisely predict aeroDynamic lift and moment coefficients in some geometric angles of attack. Moreover, unsteady effects of flow sequences in downstream of the cross section have not been included in the model similar to most of the Dynamic stall models. In order to increasing the precision of the aeroDynamic coefficients, modification and development of the boeing-vertol model is the main goal of this research witj considering the unsteady wake effects of flow sequences. Hence, unsteady aeroDynamic theory based on Wagner function was used to consider the unsteady wake effects and to introduce an effective angle of attack including airfoil degrees of freedom and their derivatives for both bending and pitching oscillations. The aeroDynamic lift coefficient of the Boeing-Vertol model was improved by using the introduced effective angle of attack and flow apparent mass effects. Also, a new pitching moment coefficient is introduced and is replaced in the model. The introduced aeroDynamic coefficients are validated and verified by experimental data and also compared with the original model. The obtained results represnt the correction of the lift coefficient in linear region of the static lift curve, improvement of maximum lift coefficient, corresponding angle of attack and improvement of moment coefficient in boeing-vertol model. Also, the results show that the proposed formulation enhances the boeing vertol model to predict moment coefficient in Dynamic condition. In addition, a parametric study is conducted to investigate the effects of reduced frequency on effective angle of attack and it is shown that while reduced frequency increases to 0. 36, unsteady wake effects on effective angle of attack of an airfoil reach to its maximum value. Moreover, for reduced frequencies upper than 0. 1, pitch axis location changes the characteristics of the effective angle of attack of the airfoil caused by unsteady effects of flow sequences.

In this research, Dynamic stall at sections near the rotor blade tip at a maximum cruise speed of the helicopter with an advanced ratio of 0.35 and cyclic pitching motion, has been studied using computational fluid Dynamics simulation. Unsteady Reynolds-averaged Navier–Stokes equations are solved using  model on a domain discretized into a hybrid mesh using finite volume discretization method. Numerical simulation is validated using experimental results of AH1-G helicopter flight tests. Comparison of results indicates that present numerical results match with experimental data well. Dynamic stall occurs as a result of a shock wave in the advancing side which affects the lift coefficient. Interestingly, the effect of the shock wave on the lift coefficient in the regions closer to the blade tip is weakened due to the tip vortex penetration. As a result, few changes are seen in the lift coefficient in these regions in comparison to those of the inner regions of the blade. In addition, the maximum value of lift coefficient in the section closer to the blade tip reduces by 10.2% in comparison to that of the most inner section. Results show that despite the formation of the leading-edge vortex, especially in the inner most sections of the blade, severe Dynamic stall does not occur in the retreating side.  In fact, this is due to the weakening of the leading edge vortex by the effect of the radial flow.

Plasma actuator is one of the newest devices in flow control techniques which can delay sepration by inducing external momentum to the boundary layer of the flow. The purpose of this paper, Dynamic stal behavour of a NACA0012 airfoil undergoing pitching motion has been studied by a numerical approach in the present and without plasma actuator. The oscillation frequency and amplitude and the Reynolds number were found to be the major contributors in Dynamic stall. The flowfield structure and the associated vortices for this airfoil as well as the impact of the oscillation frequency on aeroDynamic efficiency were also studied. The simulations were two dimensinal and the k-ω,SST turbulence model were utilized for the present analysis. The results show that in without plasma actuator increasing the oscillation frequency and amplitude, postpones the Dynamic stall to higher angles of attack. Furthermore, as increasing the Reynolds number, both the lift coefficient and the width of the associated hysteresis loop decrease. But when plasma actuator is on, Dynamic stall not happen and aeroDynamic coefficients improved. The flow field structure revealed that the main cause of the Dynamic stall is a series of low pressure vortices formed at the leading edge which shed into downstream and separate from the surface. A secondary vortex will then appear and increases the lift coefficient dramatically. But when plasma actuator is on, sepration is delay and power and size vortex much reduced.

Plasma actuator is one of the newest devices in flow control techniques which can delay sepration by inducing external momentum to the boundary layer of the flow. The purpose of this paper, Dynamic stal behavour of a NACA0012 airfoil undergoing pitching motion has been studied by a numerical approach in the present and without plasma actuator. The oscillation frequency and amplitude and the Reynolds number were found to be the major contributors in Dynamic stall. The flowfield structure and the associated vortices for this airfoil as well as the impact of the oscillation frequency on aeroDynamic efficiency were also studied. The simulations were two dimensinal and the k-ω SST turbulence model were utilized for the present analysis. The results show that in without plasma actuator increasing the oscillation frequency and amplitude, postpones the Dynamic stall to higher angles of attack. Furthermore, as increasing the Reynolds number, both the lift coefficient and the width of the associated hysteresis loop decrease. But when plasma actuator is on, Dynamic stall not happen and aeroDynamic coefficients improved. The flow field structure revealed that the main cause of the Dynamic stall is a series of low pressure vortices formed at the leading edge which shed into downstream and separate from the surface. A secondary vortex will then appear and increases the lift coefficient dramatically. But when plasma actuator is on, sepration is delay and power and size vortex much reduced.

In this study, the effect of blowing jet parameters, which include location, blowing angle, momentum coefficient, and orifice area, on the average aeroDynamic coefficients and aeroDynamic performance parameter of the NACA0012 is investigated. The airfoil has a sinusoidal oscillating motion around a quarter of the chord. As a result of this movement, the angle of attack of the airfoil changes from -5 to 25 degrees. Five locations by 1, 4, 6, 10, and 20% of the chord length were examined and the results showed that the placement of the jet at 1% of the chord length is more appropriate than the other locations and significantly improves performance and have more impact on the flow control. Three angles of 30, 60, and 90 degrees were considered as the angles of the blowing jet and it was observed that the angle of 60 degrees is better in controlling the flow than the other two angles and this effective angle decreases when orifice length increases. The results show that there is no regular uptrend or downtrend for the angle effect. For the jet momentum coefficient, which actually aimed to investigate the effect of blowing jet velocity, the values of 0.14, 0.08, and 0.04 were considered while the orifice length is considered constants for these cases. The results showed that increasing the blowing jet velocity as well as increasing the jet orifice length improves aeroDynamic performance, although increasing the orifice length at a constant momentum coefficient decreases the mean of lift coefficient.

This paper presents the investigation on the phenomenon of a deep Dynamic stall at the Reynolds number of the order of 105 over an oscillating NACA 0012 model. Wind tunnel experiments are conducted to investigate the aeroDynamic characteristics of the upstroke and downstroke phase associated with the sinusoidal pitching motion of the airfoil using the technique of surface pressure measurements and Particle Image Velocimetry. The validation of the lift curve slope of upstroke and downstroke with the Prandtl’ s thin airfoil theory reveals the fact of massive flow separation during the deep Dynamic stall regime. Numerical simulations are performed using Reynolds averaged Navier Stokes turbulence models such as RNG K-є and SST models. The data obtained from these models have been compared with the experimental data to investigate the aeroDynamic features of the deep Dynamic stall regime. The comparison shows that the URANS with K-ε model is in good agreement with the experimental data within the reasonable regime.

Dynamic stall is a phenomenon which appears due to the vortex shedding on the oscillating wing section at high angles of attack. Occurrence of the Dynamic stall causes a severe decrease in the lift force and huge increase in the drag force. The Co-Flow Jet (CFJ) is one of the active flow controls to prevent this phenomenon. In this paper, the effect of this active flow control on the NACA 0025 airfoil for different Reynolds numbers is investigated. For numerical solution of the fluid flow, the Reynolds-averaged Navier-Stokes equations in two-dimensional, incompressible, and unsteady form with the SST-k-ω turbulence model is solved using an in-house computer code. The developed code is validated with the previous experiment data, and a fairly good agreement is observed. In order to investigate the effects of the CFJ, three different momentum coefficients of 0. 05, 0. 07 and 0. 08 and five Reynolds numbers of 5×104, 7. 5×104, 105, 1. 5×105, and 3×105 are studied. It is found in the examined cases that the baseline airfoil in the Reynolds numbers of 105 and lower has different behavior compared to the higher Reynolds numbers,while in order to eliminate the Dynamic stall, it requires more jet momentum of 0. 08, while for the higher investigated Reynolds numbers, by applying the jet momentum of 0. 07, the Dynamic stall is completely eliminated.

The necessity in the analysis of Dynamic stall becomes increasingly important due to its impact on many streamlined structures such as helicopter and wind turbine rotor blades. The present paper provides Computational Fluid Dynamics (CFD) predictions of a pitching NACA 0012 airfoil at reduced frequency of 0. 1 and at small Reynolds number value of 1. 35e5. The simulations were carried out by adjusting the k − ε URANS turbulence model in order to damp the turbulence production in the near wall region. The damping factor was introduced as a function of wall distance in the buffer zone region. Parametric studies on the involving variables were conducted and the effect on the prediction capability was shown. The results were compared with available experimental data and CFD simulations using some selected two-equation turbulence models. An improvement of the lift coefficient prediction was shown even though the results still roughly mimic the experimental data. The flow development under the Dynamic stall onset was investigated with regards to the effect of the leading and trailing edge vortices. Furthermore, the characteristics of the flow at several chords length downstream the airfoil were evaluated.

Dynamic stall behavior of a NACA0012 airfoil undergoing pitching motion has been studied by a numerical approach. The turbulence intensity, oscillation frequency and amplitude and the Reynolds number were found to be the major contributors in Dynamic stall. The flow field structure and the associated vortices for this airfoil as well as the impact of the oscillation frequency on aeroDynamic efficiency were also studied. The simulations were two dimensional and the k-w SST turbulence model was utilized for the present analysis. The results show that increasing the oscillation frequency and amplitude and the turbulence intensity, postpones the Dynamic stall to higher angles of attack.Furthermore, as the Reynolds number is increased, both the lift coefficient and the width of the associated hysteresis loop decrease. The airfoil aeroDynamic efficiency variation with oscillation frequency has been shown to have a maximum point for all angles of attack considered. The flow field structure revealed that the main cause of the Dynamic stall is a series of low pressure vortices formed at the leading edge which shed into downstream and separate from the surface. A secondary vortex will then appear and increase the lift coefficient dramatically. The present simulation results are in a good agreement with those found in the literature.