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The aerodynamic loads of wind turbine blades are substantially affected by dynamic stall induced by the variations of the angle of attack of local airfoil sections. The purpose of the present study is to explore the effect of leading-edge protuberances on the fluctuation of the aerodynamic performances for wind turbine airfoil during dynamic stall. An experimental investigation is carried out by a direct force measurement technique employing force balance at a Reynolds number Re = 2 × 105. The phase-averaged and instantaneous aerodynamic loads of the pitching airfoil, including the baseline and the wavy airfoil, are presented and analyzed. The phase-averaged results indicate that the effects of dynamic stall for the wavy airfoil can be delayed or minimized compared to the baseline airfoil, and the negative damping area of the wavy airfoil is significant decreased in full-stall condition. These effects of leading-edge protuberances are more notable at a higher reduced frequency. For the instantaneous aerodynamic loads of the wavy airfoil, there is an observable reduction in fluctuations compared with baseline case. Furthermore, spectral analysis is applied to quantitatively undercover the nonstationary features of the instantaneous aerodynamic loads. It is found that the leading edge protuberances can reduce the harmonics of the aerodynamic force signal, and enhance the stability of the aerodynamic loads under different reduced frequencies. In conclusion, leading-edge protuberances are found effective to reduce the fluctuation characteristics of the aerodynamic loads during the dynamic stall process, and help to improve the stability and prolong the service life of the wind turbine blades.

Horizontal axis wind turbine (HAWT) often works under yaw due to the stochastic variation of wind direction. Yaw also can be used as one of control methods for load reduction and wake redirection of HAWT. Thus, the aerodynamic performance under yaw is very important to the design of HAWT. For further insight into the highly unsteady characteristics aerodynamics of HAWT under yaw, this paper investigates the unsteady variations of the aerodynamic performance of a small wind turbine under static yawed and yawing process with sliding grid method, as well as the there-dimensional effect on the unsteady characteristics, using unsteady Reynolds-averaged Navier–Stokes (URANS) simulations. The simulation results are validated with experimental data and blade element momentum (BEM) results. The comparisons show that the CFD results have better agreement with the experimental data than both BEM results. The wind turbine power decreases according to a cosine law with the increase of yaw angle. The torque under yaw shows lower frequency fluctuations than the non-yawed condition due to velocity component of rotation and the influence of spinner. Dynamic yawing causes larger fluctuate than static yaw, and the reason is analyzed. The aerodynamic fluctuation becomes more prominent in the retreating side than that in the advancing side for dynamic yawing case. Variations of effective angle of attack and aerodynamic forces along the blade span are analyzed. The biggest loading position moves from middle span to outer span with the increase of yaw angle. Three-dimensional stall effect presents load fluctuations at the inner board of blade, and becomes stronger with the increase of yaw angle.

Insufficient understanding of flow instability interaction between hump characteristic and the rotating stall is a major problem in the application of mixed flow pumps. In this paper, the unsteady flow characteristics of a mixed flow pump are studied numerically to understand the hump characteristic generating mechanism using a modified SST k-ω partially averaged Navier-Stokes (MSST PANS) model. The predicted pump characteristic curves show good agreement with experimental data. Fundamental results on time-averaged flow pattern and energy loss distributions depict that the substantial rise of energy loss in the pump impeller is the main reason for the hump characteristic. Detailed flow structures show the peak regions of circumferential vorticity (Ωθ) and boundary vortex flux (BVF) are located at the blade tip near the trailing edge (TE) and leading edge (LE). Further analysis on the flow angle and blade loading are carried out. The results indicate that the peak region located near the LE of spanwise = 0.8 and streamwise = 0.3 is the source of the rotating stall evolution. In this process, the substantial reduction in the blade loading occurs near the LE, eventually induces the head drop. Finally, pressure fluctuations analysis depicts that low-frequency signal is observed induced by the rotating stall evolution.

A highly loaded, high-speed turn flight of Airbus Helicopters' Bluecopter demonstrator helicopter is simulated to investigate dynamic stall using a loose computational fluid dynamics/structural dynamics (CFD/CSD) coupling of the flow solver FLOWer and the rotorcraft comprehensive code CAMRAD II. The rotor aerodynamics is computed using a high-fidelity delayed detached-eddy simulation (DDES). A three-degree-of-freedom trim of an isolated rotor is performed, yielding main-rotor control angles that agree well with the flight-test measurements. The flow field in this flight condition is found to be highly unsteady and complex, featuring massively separated flow, blade–vortex interaction, multiple dynamic-stall events, and shock-induced separation. The computed pitch-link loads are compared to flight-test measurements. This shows that all CFD/CSD cases underpredict the amplitudes of the flight test and yield phase shifts. However, overall trends agree reasonably. Also, varying the computational setup reveals that the shear stress transport–DDES turbulence model performs better than Spalart–Allmaras–DDES, that the consideration of the rotor hub and fuselage improves the agreement with flight-test data, and that the elastic twist plays only a minor role in the dynamic-stall events.

The delayed detached-eddy simulation with adaptive coefficient (DDES-AC) method is used to simulate the baseline and leading-edge undulation control of dynamic stall for the reverse flow past a finite-span wing with NACA0012 airfoil. The numerical results of the baseline configuration are compared with available measurements. DDES and DDES-AC perform differently when predicting the primary and secondary dynamic stalls. Overall, DDES-AC performs better owing to the decrease of grey area between the strong shear layer and the fully three-dimensional separated flow. Moreover, the effects of the undulating leading-edge on the forces, lift gradients, and instantaneous flow structures are explored. Compared with the uncontrolled case, the lift gradient in the primary dynamic stall is reduced from 18.4 to 8.5, and the secondary dynamic stall disappears. Therefore, periodic unsteady air-loads are also reduced. Additionally, the control mechanism of the wavy leading edge (WLE) is also investigated by comparison with the straight leading edge (SLE). No sudden breakdown of strong vortices is the main cause for WLE control.

In this study, viscous, turbulent, and steady flow around an airfoil near the water surface has been simulated through a numerical method. In this simulation, Navier-Stokes equations have been solved using the finite volume method with a discretized second-order accuracy and PIMPLE algorithm. The Volume of Fraction (VOF) method has been employed to predict the free surface flow. A part of the simulation results has been validated through numerical and experimental data. Besides considering the style of flow separation in the angles of numerous attacks and airfoil static stall near the surface of the water. For this purpose, the airfoil simulation has been processed airfoil in the 68,000 Reynolds number, angle of attack of 2.5 to 11 degree and different distances from the water surface ( h/c = 0.5, 1, ). In a larger angle of attacks, flow is initially separated from the leading edge of the surface, and then it attaches to the surface at a lower point. This reattachment leads to an increase in adverse pressure gradient and the formation of a larger separation in the downstream of the airfoil. The pressure gradient dramatically increases, and the flow gets separated from the upstream of the airfoil. Upon lowering distance from the surface, static stall takes place at a higher point and a lower angle of attack, respectively.

The enormous loss of momentum leads to stall and adversely affects the aerodynamic performance of aeroplane wings which may lead to a disaster, more importantly, risking the safety of the aeroplane by putting lives of passengers on it in danger. Therefore, this paper focuses on the enhancement of aerodynamic characteristics of NACA 23012 through the mitigation of flow separation and delay of the stall at higher angles of attack by using suction for Reynolds number (Re) = 3.4 x 106 . Considering the different suction features such as suction width, suction position, and suction coefficient, the separation delay capability of a suction control is studied. Also, the lift to drag ratio and the impact of energy consumption variation during the control technique are used for estimating the control effects. The Reynolds Average Navier-Stokes (RANS) equations are employed together with the Menter’s shear stress turbulent model. The result of this study revealed that the jet position just behind the separation point at 0.2 % of the chord length shows an outstanding control outcome on the separation and stall, thereby increasing the lift. The lift to drag ration increased proportionately when the suction jet coefficient was increased. At suction coefficient Cq = 0.00225, a 92.1% drag reduction and 72.7% lift enhancement is observed. Hence, the stall angle is moved beyond 21.5o from an initial angle of 16° and the more energy was saved at a high angle of attack.

Purpose The purpose of this study is to the modeling of the dielectric barrier discharge (DBD) actuator on the Eppler 387 (E387) airfoil in low Reynolds number conditions. Design/methodology/approach A validated direct-forcing immersed boundary method is used to solve the governing equations. A linear electric field model is used to simulate the DBD actuator. A ray-casting technique is used to define the geometry. Findings The purposed model is validated against the former studies. Next, the drag and lift coefficients in the static stall of the E387 airfoil are investigated. Results show that when the DBD actuator is on, both of the coefficients are increased. The effects of the location, applied voltage and applied frequency are also studied and find that the leading-edge actuator with higher voltage and frequency has better improvement in the forces. Finally, the dynamic stall of the E387 with the DBD actuator is considered. The simulation shows that generally when the DBD is on, the lift coefficient in the pitch-up section has lower values and in the pitch-down has higher values than the DBD off mode. Practical implications It is demonstrated that using the DBD actuator on E387 in the low Reynolds number condition can increase the lift and drag forces. Therefore, the application of the airfoil must be considered. Originality/value The results show that sometimes the DBD actuator has different effects on E387 airfoil in low Reynolds number mode than the general understanding of this tool.

The use of dielectric barrier discharge plasma actuators to suppress traveling stall cells and thereby improve stall margin is examined. This involved the design and construction of a nonconductive...

Large-eddy simulations of the onset of dynamic stall for a NACA 0012 airfoil, pitching at a constant rate, are performed for a chord Reynolds number of 1.0×106. Using four simulations, spanning M∞=...

Stalling of residential air conditioners (RACs) following a transient fault can delay the voltage recovery ranging from 3 to 20 seconds. This phenomenon is referred to as fault-induced delayed voltage recovery (FIDVR). Several aggregated RAC models have been presented to replicate the actual FIDVR events. However, they usually have two main drawbacks as: i) considering independent stalling voltage and stalling time for all RACs, ii) resulting in only two ultimate modes; either 100% stalled or 100% normal running. This paper amends the abovementioned shortcomings by proposing a simple and effective performance-based RAC model. The proposed model utilizes a simple explicit relation between the stalling voltage and its required time to stall the RAC, furthermore, it can stall a fraction of RACs based on the voltage dip characteristics. The experimental results support the proposed model and computer simulations employing the IEEE 118-bus test system and Iran's electricity grid reveal the effectiveness of the proposed model.

A centrifugal compressor is a typical compressed air energy storage device. In order to ensure the safety of the compressed energy storage process in the compressor, the internal unsteady flow phenomena need to be closely monitored, especially some serious ones like stall and surge. It is necessary to explore the mechanism of flow instabilities under different conditions. A centrifugal air compressor was tested with a vaneless diffuser and a variable vaned diffuser with five different vane setting angles, respectively. Various diffuser types resulted in various modes of flow instabilities prior to surge. The vaneless region between the impeller and the diffuser was focused on. Multiple high-speed sensors were arranged along the circumferential direction. The pressure signals at all these positions were being measured and collected in real time as the compressor was slowly throttled into surge. This paper emphasizes on the influence of matching between the impeller and the diffuser on the flow instability. The experimental results showed that the diffuser vane setting angle affected the stall characteristics. Due to the asymmetry of the volute, the circumferential pressure distribution was always severely distorted prior to surge. A high-pressure region appeared near the volute tongue, and a low-pressure region was formed away from the volute tongue. In the case of the vaned diffuser with non-design installation angle and the vaneless diffuser, the rotating stall signal was originated in the low-pressure region and propagated circumferentially. However, in the case of the vaned diffuser with the design installation angle, the circumferential high-pressure region became the most sensitive region for the generation of stall, and another form of instability occurred there. Both the inducement and development of these flow instabilities have been studied. The dynamic experimental research on the compressor matching different types of diffusers could be a good case supplement.

Voluminous research exists on the effect of combination of low Reynolds number flow over low aspect ratio flat plate wings for usage in micro-aerial vehicles and UAVs, as they also fly under similar conditions and combination of low aspect ratio and low Reynolds number. It has been well established by many researchers (Torres and Mueller in AIAA 42(5):865–873, 2004; Okamoto and Azuma in AIAA 49(6):1135–1150, 2011; Mizoguchi and Itoh in AIAA 51(7):7–52, 2013; Shields and Mohseni in AIAA 50(1):85–99, 2012; Mizoguchi and Kajikawa in Trans Jpn Soc Aero Sp Sci 59(2):56–63, 2016; Pelletier and Mueller in J Aircr 37(5):825–832, 2000; Liu and Hsiao in J Mech 28(01):77–89, 2012; Taira and Colonius in J Fluid Mech 623:187–207, 2009) that low aspect ratio wings in low Reynolds number flow gives a higher lift coefficient and stall angle is also increased which is quite different from the high aspect ratio wing. Also, it is a well-established fact in the existing literature that linear theories cannot be applied as in case of high aspect ratio wings because of non-linear increase in lift coefficient. The reason for such increase in lift and stall angle has been attributed to wingtip vortices that become prominent in low aspect ratio wings called the vortical lift which keeps on increasing up to a certain high angle of attack till it interacts with the separation bubble. In the past, very limited work has been done to study the vortices and their contribution in enhancement of aerodynamic characteristics. In the present work, flat plate model with corrugations at an aspect ratio (AR) of 1.0 are studied and compared with flat plate wing of AR 1.0; also flow visualization in water tunnel has been carried out at Reynolds number of 102 to ascertain the flow physics involved.

A massive parallel aeroelastic simulation platform has been built to investigate the first 1.5-stage fan of an aeroengine at rotating stall. The Computational Fluid Dynamics (CFD) solver and Computational Structural Dynamics (CSD) solver are coupled directly by non-matching mesh interfaces. The unsteady rotor/stator interaction is solved by the Sliding Mesh Interface method. The original rotor blades are shrouded by the midspan shrouds. An unshrouded fan is also created to investigate the effects of the midspan shrouds. Both the shrouded fan and unshrouded fan have stable aeroelasticity at the designed state. At rotating stall, the stalled region rotates at 30% of the rotor speed on the absolute reference frame. The energy spectrum of the rotating stalled flow is measured quantitatively. It shows that the first two order excitations are much stronger than the higher order excitations. In the flow of rotating stall, the fifth backward travelling wave mode of shrouded fan is resonated by the fifth excitation of the rotational stalled flow because the rotational speed of the stalled region coincides with the modal rotational speed, while for the unshrouded fan, the first bending mode is resonated by the second excitation of the rotational stalled flow, forming two waves in the circumference of the annulus blades. At rotating stall, the vibration of the shrouded blades is still under control but the vibration of the unshrouded blades is diverged and out of control. A novel tool, i.e., resonance map, is proposed to predict the resonance. It provides a perspective to explain the effects of midspan shrouds at a theoretical level, and it would also be helpful in the structural design of blades. Keywords: Aeroelasticity, Computational Fluid Dynamics, Midspan shrouds, Rotating stall, Turbomachine blades

Have stakeholders increased their influence over firm behavior in the digital age? We draw from cognitive theory to argue that although social media has made it easier for stakeholders to broadcast...

Accurate predictions of the cycle-to-cycle variations in aerodynamic loads on wind turbine blades are important for estimating the life cycle costs of new wind power plants. Physical modelling of wind turbine systems is usually based on the average values of the aerodynamic parameters in the design process which would eliminate some important physical phenomena associated with the unsteady flows over the rotating blades. This paper presents a new algorithm implemented in a Blade Element Momentum (BEM) model to predict the cycle-to-cycle variations in the aerodynamic loads over multiple rotor rotations (cycles) for yawed rotors operating in natural flow conditions. In this approach, simulation theory developed by the Pacific Northwest Laboratories (PNL) is used to generate input turbulent wind speed time series on the whole rotor disc of two three-bladed downwind horizontal axis wind turbine ( NREL Phase II and III ) rotors. The unsteady influences of the dynamic stall kinematics and tower shadow are also considered in the current algorithm along with the phenomenon of stall delay. Predictions from the unsteady BEM model are compared to measured results and conclusions are drawn about the ability of the model to accurately simulate blade response to turbulence.

The multiscale development of dynamic stall was studied using a scale-based modal analysis technique. Time-resolved velocity field data around an airfoil during dynamic stall were used for this ana...

Morphing wings can improve aircraft performance during different flight phases. Recently research has focused on steady aerodynamic characteristics of the morphing wing with a flexible trailing-edge, and the unsteady aerodynamic and stall characteristics in the deflection process of the morphing wing are worthy further investigation. The effects of the angle of attack and deflection rate on aerodynamic characteristics were examined, and based on the aerodynamic characteristics of the morphing wing, a method was developed to delay stall by using the flexible periodic trailing-edge deflection. The numerical results show that the lift coefficients in the deflection process are smaller than those in the static situation at small angles of attack, and that the higher the deflection rate is, the smaller the lift coefficients will be. On the contrary, at large angles of attack, the lift coefficients are higher than those in the static case, and they become larger with the increase of the deflection rate. Further, the periodic deflection of the flexible trailing-edge with a small deflection amplitude and high deflection rate can increase lift coefficients at the critical stall angle. Keywords: Aerodynamics, Computational fluid dynamics, Flexible structures, Morphing wings, Stall angle

The internal blown flap was numerically simulated. Firstly, a parameterization method was developed, which can properly describe the shape of the internal blown flap according to such geometrical parameters as flap chord length, flap deflection, height of blowing slot and its position. Then the reliability of the numerical simulation was validated through comparing the pressure distribution of the CC020-010EJ fundamental generic circulation control airfoil with the computational results and available experiment results. The effects of the geometrical parameters on the aerodynamic performance of the internal blown flap was investigated. The investigation results show that the lift coefficient increases with the increase of flap chord length and flap deflection angle and with the decrease of height of blowing slot and its front position. Lastly, a method of optimal design of the geometrical parameters of the internal blown flap was developed. The design variables include flap chord length, flap deflection, height of blowing slot and its position. The optimal design is based on maximum lift coefficient, the angle of attack of 5 degrees and the design constraint of stall angle of attack of less than 9 degrees. The optimization results show that the optimal design method can apparently raise the lift coefficient of an internal blown flap up to 1.7.

Experimental analyses of synthetic jet control (SJC) effects on aerodynamic characteristics of rotor in steady state and in hover were conducted. To ensure the structural strength of rotor and enough interior space for holding the synthetic jet actuators (SJAs), a particular blade with a frame-covering structure was designed and processed, and the experiment was conducted with low free stream velocities and rotor rotation speeds. There were three test conditions. In steady state, there were three free stream velocities (10m/s, 15m/s and 20m/s). In hover state, the rotor was worked with two rotation speeds of 180RPM and 240RPM. In forward flight, the rotor was worked with a rotation speed of 180RPM and a free stream velocity of 7.5m/s. To measure the synthetic jet control effect on rotor in stall, the range of collective pitch was set from 10° to 28° in steady state. The aerodynamic forces and sectional velocity field were measured by using the six-component balance and the Particle Image Velocimetry (PIV) system in the wind tunnel. Flow control effects on the blade based on the synthetic jets (SJ) were experimentally investigated with different jet parameters, such as jet locations, jet angles, and jet velocities. In steady state, the jet closer to the leading edge, and the jet angle of 90° had more advantages in improving the aerodynamic characteristics. Furthermore, the aerodynamic forces and sectional velocity field measurement of rotor in hover were conducted, it showed that SJAs could increase flow velocity at the upper surface, which led to lower upper surface pressure. As a result, the normal forces of rotor with two rotation speeds were increased significantly. These results indicated that the synthetic jet has a capability of increasing the normal force and delaying or preventing the stall of rotor.

Wind turbines in complex environments could encounter dynamic loads that deviate from design values, while the dynamic stall problem could reduce the reliability and safety of wind turbines. In order to accurately evaluate the aerodynamic characteristics of wind turbine blades, the influence on the dynamic stall characteristics of wind turbine airfoil under complex wind farm environmental factors is studied. The variation law of lift coefficient under different working conditions is obtained, and the pressure coefficient and flow field distribution with times are given. The evolution characteristics of vortex on airfoil surface are analyzed. The results show that when the unsteady effect is significant, the closed-loop area of the blades lift coefficient is gradually larger, and the angle of attack for the maximum lift is proportional to it; when the oscillation angle becomes larger, the upper and lower extreme values of the lift force increase, and the lift curve is more obvious in the shape of the oblique “$$\infty$$.” When the blade appears to be in dynamic stall, the lift coefficient increases quickly and has a hysteresis effect. The maximum increase is up to 46.15%, and the hysteresis effect occurs. The instability of the shear layer, the formation and quenching of the trailing edge vortex are the main causes of the hysteresis effect for the blade lift coefficient. Reasonable evaluation of its value is of great significance for wind turbine blade design in complex wind farm environment.

At present, off-design centrifugal compressor suffers from great instabilities, such as rotating stall and surge, which lead to work failure, strong pressure fluctuations, and violent vibrations of the entire system. One cause of these instabilities is the rotating stall occurring in vaneless diffuser. In this review, recent researches on rotating stall of vaneless diffuser in centrifugal compressor are reviewed, and its inductive mechanism and characteristic description are emphatically introduced. In vaneless diffuser, the rotating stall starts up and fully develops under non-design conditions with the characteristic of increasing pressure fluctuation amplitude. The negative effects of rotating stall include reverse flow and vortex, both of which have significant effects on the instability of centrifugal compressors. Some improvement methods of rotating stall are also introduced, and a lot of examples are given. Finally, some suggestions and discussions are put forward for the future work.

A one GPa controlled clearance piston gauge used as a high pressure standard at KRISS was investigated. In establishing the high pressure standard, the distortion coefficient is one of the most important parameters, but it is not easy to determine. The conventional analysis using a Heydemann–Welch model has a problem when determining a stall pressure, because of the nonlinear characteristic of fall rates with respect to pressure. Some metrological characteristics, such as the distortion coefficient of the piston–cylinder assembly in free deformation (FD) mode and the jacket distortion coefficient and stall pressure in controlled clearance (CC) mode, were investigated using a finite element analysis (FEA). In particular, it was determined that the relation of cubic fall rate to jacket pressure becomes nonlinear above 600 MPa. The FEA results were verified by comparison with a 500 MPa pressure standard and a fall rate measurement of 1 GPa. The most important parameter, the distortion coefficient in the FD and CC mode, was determined to be (7.59 ± 0.24) × 10−7 MPa−1 and (3.35 ± 0.39) × 10−7 MPa−1, respectively. A zero-distortion coefficient was obtained in the FEA when around 17% of the applied pressure was used as the jacket pressure. This value was similar to the experimentally determined result of around 20%.

To study the stall mechanism in a wide vaneless diffuser, a three-dimensional centrifugal compressor model was established. The numerical simulation method was used to study the fluid characteristics in a centrifugal compressor with a wide vaneless diffuser under stall conditions. The simulated results showed that the wide vaneless diffuser stall was related to the instability of the core flow. Due to the influence of the reflow, the core flows moved between the hub and shroud side and the directions were not consistent. This was called core flow distortion, which triggered the rotating stall. The reflow first appeared at the outlet boundary of the diffuser, and the reflux gradually expanded to the inside of the diffuser as the impeller rotated. Eventually, the reflux zones grew and merged, forming a stable low-speed fluid regiment, which produced a stall cell. The stall cells rotated around themselves with an opposite direction with impeller.

China’s modern education system urgently requires entrepreneurship and innovation education. By reading literature, launching questionnaire, and giving interviews, we found that most colleges and universities in China neglect developing students’ practical ability in entrepreneurship and innovation education, and that the proposal made in existing research to add practical training just focuses on macro-guidance, with no specific solutions given. Therefore, we innovatively propose to introduce street stall economy into the curriculum based on the features and feasibility of China’s street stall economy, and offer recommendations for management and assessment by drawing on the experience of other places in administering street stall economy.

The unsteady flows past a NACA0015 airfoil near stall at a high Reynolds number are investigated numerically using the delayed detached-eddy simulation with adaptive coefficient model. The unsteady...

A novel high-frequency flow control strategy for mitigation of dynamic stall is demonstrated for the case of a finite wing using high-fidelity wall-resolved implicit large-eddy simulations. A NACA ...

In this study, we experimentally investigate the effects of mainstream turbulence intensity (Ti) on a leading-edge separation bubble under low-Reynolds number (Rec) conditions range of 2.0 × 104 to 6.0 × 104. We used a flat plate to fix a separation point at the leading edge. Also, we visualized the behavior of the leading-edge separation bubble using the smoke wire technique and Particle Image Velocimetry (PIV) measurement. Furthermore, we measured the effect of Ti on the turbulent transition process in the separated shear layer using a hot-wire anemometer. The results indicate that the bypass transition for large Ti causes the turbulent transition, and so accelerates the reattachment of the separated shear layer. The results show that the bypass transition promotes the reattachment of the separated shear layer to maintain the leading-edge separation bubble on the upper surface even at high angles of attack, increasing the stall angle.

In wind turbines, increased blade surface roughness accelerates fluid separation and reduces blade aerodynamic performance. As a flow separation control device, a vortex generator can be used to modulate the fluid flow in the boundary layer and reduce the adverse effect of increased roughness on the aerodynamic performance of the blade. To improve the effect of the vortex generator, a bionic vortex generator based on the anatomy of manta rays is here proposed and commercial CFD software ANSYS CFX is used for numerical simulation. We found that when the equivalent roughness is 0.045 mm, the starting performance of the DU97-W-300 airfoil is consistent with the experimental value observed for installed zigzag tape. Under different roughness conditions, the optimization of the blade aerodynamic performance of the bionic vortex generator is up to 10% higher than that of the triangular vortex generator. The presence of the vortex generator not only increases the maximum lift coefficient of the contaminated blade, but also delays the stall angle, and at the same time increases the suction power of the entire suction surface of the blade. The vortex generator has little effect on the pressure surface.

In the past extensive research has been carried out, to study the effect of Gurney flap (GF) on symmetric and cambered airfoil for its usage in low Reynolds number regime. Use of GF at the trailing edge of the airfoil enhances the lift due to increase in the effective camber of the airfoil, which in turn improves the aerodynamic efficiency i.e. Cl/Cd. In the present study, Eppler 423 airfoil is used to first understand the aerodynamics of such a highly cambered airfoil and later GF of various sizes were added on it to understand the change in flow dynamics achieved by adding the GF and their impact on aerodynamic parameters such as Cl, Cd and Cl/Cd. Eppler 423 being a highly cambered airfoil produces high lift coefficient and smoother stall and by adding the GF of various sizes the performance of Eppler 423 improves tremendously and reason for this enhanced performance and effect of size of GF are presented in this paper. Vortex Generators (VG) generate counter rotating vortices that allow the flow to remain attached even at high angles of attack. Also, effect of adding VG at the leading edge of Eppler 423 aerofoil is presented in this paper. At last, results obtained from combination of VG at leading edge and GF at trailing edge on Eppler 423 aerofoil are discussed at length.

This paper presents the design of an automatic deep stall recovery algorithm for large transport aircraft using optimal trajectory planning. Deep stall is a condition where an aircraft is trapped in a nose-high stall condition and its elevators cannot produce enough nose-down pitching moment to recover the aircraft from the stall. The NASA Generic Transport Model (GTM) is used as the basis for the design and verification of the system. The aerodynamic model of the NASA GTM simulation model is modified to exhibit deep stall behaviour. Simulations are performed to show that the modified aircraft model can be pushed into deep stall, and cannot be recovered using elevator actions only. The deep stall recovery task is formulated as an optimal path planning problem and solved using an A* search algorithm to find the optimal sequence of control actions and the resulting optimal state trajectory to escape from the deep stall. The A* algorithm performs the planning using a simplified, three-degrees-of-freedom (3DOF) aircraft model that models only the fast rotational dynamics. The automatic deep stall recovery is then verified in simulation using the full six-degrees-of-freedom (6DOF) NASA GTM aircraft model. The simulation results show that the system successfully recovers the aircraft from deep stall. The optimal sequence of control actions first uses the rudder to yaw the horizontal tailplane out of the aircraft’s own wake to regain elevator effectiveness, and then uses the elevators to pitch the nose down and recover from the stall.

There are multiple approaches of design for Vertical Axis Wind Turbines (VAWT) that have been studied by engineers and leaps have been made in high performing innovations. By harnessing the energy from these wind turbines, the problem of roadside lights shortage can be solved. This can help to prevent the accidents while providing clean energy. The importance of coastal areas like Australian beaches regarding wind turbines cannot be neglected as a higher number of people like to live near coastal vicinity. Also, most of the freeways in Australia expand across the sea. In this paper, one such design has been analyzed to implement across the highways. But still with many advancements in technology, an immense gap is present in the research of implementation of VAWTs. The design discussed in the current study is a VAWT which can be installed on the side of the highway roads to provide clean and cheap energy for illuminating the roads. Computational Fluid Dynamics (CFD) was conducted on the blades of the turbine to analyze its performance under operating conditions. Furthermore, the paper elaborates the generation of drag and lift on the blades of the turbine. A wind speed of 60 km/h just produced 6.1 N force on the turbine blades as a result of drag. The cost analysis showed the cheap production of such mechanism that can provide longer service when installed.

Winged drones that fly in close proximity to obstacles or that are capable of aerial and terrestrial locomotion can benefit from protective systems that prevent damage to delicate aerial structures. Existing protective solutions focus on multi-copter drones and consist of adding structures, such as cages, mechanisms and instruments that add weight and drag. Here we describe a protective strategy for winged drones that mitigates the added weight and drag by means of increased lift generation and stall delay at high angles of attack. The proposed structure is inspired by the wing system found in beetles and consists of adding an additional set of retractable wings, named elytra , which can rapidly encapsulate the main folding wings when protection is needed.

The influence of Reynolds number, roughness and turbulence on the onset of wall stall is up to now not sufficiently understood. To shed some light onto the interdependency of near wall flow with growing swirl component, the simplest “machine” is tested. The apparatus we examine is a circular pipe at rest followed by a rotating co-axial pipe segment. In the sense of a generic experiment this machine represents a very basic model of the inlet of an axial machine. Due to the wall shear stress a swirl boundary layer is formed in the rotating pipe segment, interacting with the axial boundary layer. The evolution of the swirl velocity profile with increasing axial distance from the rotating pipe inlet is measured for various Reynolds numbers, flow numbers and degrees of turbulence by means of Laser Doppler Anemometry. We observe a self-similarity in the swirl velocity profile, for subcritical flow number and develop a scaling law for the velocity distribution in the transition section of a rotating pipe. At critical flow number the boundary layer is separating, resulting in a ring vortex at the inlet of the rotating pipe. Our work fills the gap of previous experimental works, with respect to high Reynolds numbers and low flow numbers. The parameter field we examine is most relevant for turbomachinery application and wall stall. In addition our boundary layer resolution is sufficient to resolve the swirl boundary layer thickness. Only this high resolution enables us to generalize the experimental findings by means of a similarity distribution of the velocity profile within the swirl boundary layer.

The tip leakage flow over the blades of an axial compressor rotor adversely affects the axial rotor efficiency and can determine the onset of tip leakage stall. The performance of a new casing treatment concept in the shape of an axisymmetric recirculation channel is explored by steady Reynolds-Averaged Navier&ndash, Stokes (RANS) realizable k-&epsilon, modelling on the NASA Rotor 37 test case. The modelling exposed a number of attractive features. The casing treatment increased the stall margin at no penalty to the rotor isentropic efficiency over the rotor operating line. A recirculation in the casing channel self-activated and self-adjusted with the rotor loading to provide more passive flow control at higher rotor loading conditions. The nozzle-shaped recirculation channel outflow opposed the tip leakage jet, re-located the casing surface flow interface further downstream, and reduced the rotor blade tip incidence angle. This combination of features makes the new casing treatment particularly attractive for applications to high thrust-to-weight ratio engines, typical of high-performance jet aircraft.

A number of methods to size the wing of a new GA aircraft are presented. The chapter begins by the presentation of handy formulation to calculate various properties of trapezoidal surfaces, which is the most common planform shape for lifting surfaces like wings and horizontal and vertical tails. This is followed by the introduction of various concepts, topics, and methods for laying out the wing planform. These include aspect ratio, taper ratio, washout, and wing incidence angle. In order to help the designer evaluate the pros and cons of typical wing planform shapes, suitable for the new aircraft, a detailed discussion follows. In addition to the large number of different geometries presented detailed information about the spanwise distribution of section lift coefficients is provided to help the designer realize their properties. Next, a number of methods to evaluate the lift and pitching moment characteristics of the wing as a 3-dimensional lifting body are provided. Then, a number of issues that have to do with wing stall characteristics and how to improve them are presented. Finally, in order to help the designer in realizing the properties of the selected wing planform and airfoils, a practical version of Prandtl’s Lifting Line Theory is presented. Despite being old, this is a sophisticated numerical method used to estimate the aerodynamic properties of the wing. A computer function, written in Visual Basic for Applications, intended for use with Microsoft Excel is presented as well, allowing the reader to get to work immediately.

Milking stall dimensions have not been adapted to the increase in cow body size caused by selection for better milking performance over the past decades. Improper milking stall dimensions might limit cow comfort, could lead to stress responses during milking and thus could negatively affect cow welfare. A crossover study was conducted in an experimental milking parlor that was converted from a herringbone (HB) to a side-by-side (SBS) parlor. The milking stall dimensions were modified in length and width and for HB also in depth (perpendicular distance between rump rail and breast rail). The stall dimensions applied during the experiments ranged from much smaller than common in European dairy farming to much larger. Treatments were applied for 2 wk per milking parlor type. In each milking parlor type, a total of 30 cows, kept in 2 groups were observed during milking for behavioral and physiological stress responses and for milking performance. In addition, milk cortisol levels and somatic cell counts were measured at the end of the 2-wk period. Outcome variables were selected based on a principal component analysis and analyzed using mixed effects models reflecting the experimental design. The results showed that the first cow per milking batch required more time (on average >40 s) to enter very small HB stalls than to enter small, large and very large stalls (, Journal of Dairy Science, 105 (1), ISSN:0022-0302, ISSN:1525-3198

This paper presents the effects of airflow bleed from the rotor shroud surface between leading and trailing edges on aerodynamic and structural performances of a single-stage transonic axial compressor, NASA stage 37, using three-dimensional Reynolds-averaged Navier–Stokes equations with the k–e turbulence model. A small airflow mass flow rate is bleed throughout a rotor-bleeding ejector which designed by seven parameters: bleeding angle (°) and ejection angles (°), ejection depth (D), bleeding thickness (H), bleeding position from rotor leading edge (L) in flow direction, bleeding ejection curvature (R), and bleeding width contacted on rotor shroud surface (W). The numerical results for aerodynamic performance: total pressure ratio, adiabatic efficiency, and stall margin of a transonic axial compressor were validated with a smooth casing experimental data. A parametric study of seven design parameters of rotor-bleeding ejector above combined with a small ejection mass flow rate in a single-stage transonic axial compressor for aerodynamic and structural performances was performed. The numerical results show that all aerodynamic performance increases with bleeding airflow from rotor shroud surface, the total deformation on rotor tip leading edge in spanwise direction reduces with a very small increasing in Von-Mises stress in a reference-bleeding airflow as compared to the results of smooth casing.

Contra-rotating fan is a concept that can possibly replace the present-day conventional fans due to its several aerodynamic advantages. It has the potential to improve the stability limit and can achieve a higher pressure ratio per stage. One of the advantages of a contra-rotating fan is its capability to operate both the rotors at different speeds. In the present study, experiments are carried out at different speed combinations of the rotors and the stall inception phenomenon is captured using high-response unsteady pressure sensors placed on the casing upstream of the leading edge of rotor-1. The unsteady pressure data are investigated using wavelet and Fourier analysis techniques. It is observed that the mechanism of stall inception is different for different speed combinations. The pre-stall disturbances fall in different frequency ranges for different speed combinations. For the range of speed combinations investigated, the frequency of appearance of stall cells of rotor-1 does not depend on the speed of rotor-2. A higher speed of rotation of rotor-1 leads to a higher frequency of appearance of stall cells and a lower speed of rotation of rotor-1 leads to a lower frequency of appearance of stall cells. For all the speed combinations, there is a range of frequency where no disturbance is observed and this range is termed as the ‘no-disturbance zone’. Disturbances are observed at lower frequencies and at frequencies close to the blade passing frequency. In order to understand the flow physics in detail, computational analysis is carried out for different speed combinations of the rotors. For a higher speed of rotor-2, it is observed that the suction effect of rotor-2 is significant enough to pull the tip-leakage flow towards the axial direction. Thus, the suction effect of rotor-2 plays a significant role in determining the stall of the stage.

Shallow hydrofoils are known to produce lower lift in normal operating conditions in comparison with deep hydrofoils. However, the maximum lift capability of shallow hydrofoils at moderate speeds, which is important for transitional regimes of hydrofoil boats, is studied insufficiently. In this work, two-dimensional flow around a high-lift hydrofoil at a moderate Froude number is numerically simulated in a broad range attack angles up to the stall occurrence in both single-phase fluid and in the vicinity of free surface. It is found that nearly the same maximum lift coefficient can be produced by the shallow foil in the modeled condition as by the deep foil, but much higher attack angle is required near the free surface, which also results in larger drag. Additionally, it is shown that higher Reynolds numbers lead to higher lift coefficients, especially at large attack angles.

This paper dwells upon investigating the effect of aspect ratio (AR) variation on the aerodynamic performance of unconventional control surfaces called grid fins by virtue of a series of subsonic experiments on a simplified grid fin variant called the cascade fin. Wind tunnel tests were performed for differentAR(variable span) grid fins. The same had been investigated for different gap-to-chord ratio (g/c) variants. Results demonstrated a tangible increase in the aerodynamic efficiency as well as stall angle reduction for higherAR. Moreover, higherARleads to increased pitching moment, which emphasizes elevated hinge moment requirements. The study ensued the presence of higher deviation between the lowARfins, that is$ARcompared to the pertinent deviations between the highARfins, that is$AR\geq2$. The effect associated with these variations was termed as span effect in this paper. It was established that, the deviations arising due to this phenomena were lesser for higherg/cand higherAR. The analysis ofARvariation for differentg/cpresented a limiting value ofARreduction for stall performance enhancement. Thus, optimised selection of theg/candARvalues can lead to enhanced aerodynamic efficiency alongside an improved stalling characteristic.

Purpose 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. Design/methodology/approach 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. Findings 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. Originality/value 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.