Increment In Lift Coefficient Research Articles

Review on Vortex Generators Mounted on Airfoil to Improve the Aerodynamic Performance

In aircrafts lift is the most important factor which depends on wing performance. But during the cruising flight when the angle of attack increases there is change in the aerodynamic factors such as lift and drag forces due to the boundary layer (BL) separation. To delay or avoid this boundary layer separation and improve the aerodynamic performance of the wing vortex generators were introduced. Vortex generators helps to keep the boundary layer attached to the wings upper surface. In previous decades much study is done to control this BL on the different types of airfoils i.e. symmetrical or unsymmetrical. This review is based on study related to improve these parameters so that there will be increment in lift coefficient and reduction in drag coefficient. To overcome these parameters appropriate size, shape, length, position and incidence angle, etc should be investigated to find out the suitable one. These investigations can be performed as numerical based using software simulation and also by experimental approach in a wind tunnel. The main practical advantage of these vortex generators is to reduce parasite drag on airfoil and increase the lift force, they are also demonstrated to be critically important in other applications also.

Comparative Study of Dynamic Stall between an Aircraft Airfoil and a Wind Turbine Airfoil in an Air–Particle Flow

Dynamic stall in clean air flow has been well studied, but its exploration in air–particle (air–raindrop or air–sand) flow is still lacking. The aerodynamic performance loss of aircraft (NACA0012) and wind turbine (S809) airfoils and their differences during the hysteresis loop at different pitching parameters are also poorly understood. As shown in this paper, the reduced frequency has little effect on the value of the maximum lift coefficient increment caused by particles, but a larger one can enhance the hysteresis effect and drag the angle of attack, at which the maximum increment is obtained, from the up stroke to the down stroke. The large lift coefficient increments of two airfoils and their difference also have a similar change trend with the reduced frequency. Compared to that of NACA0012 airfoil, the increments of S809 airfoil are obviously greater at three mean angles of attack, especially at 8°, which is the commonly used operating angle. In addition, the angle of attack, at which the maximum lift coefficient is obtained, can be significantly changed by particles in two regions: one is under the effect of deep stall, the other is under the effect of light stall at a low, reduced frequency.

Optimization of the Slotted Gurney-Flap geometry applied to NACA 0012 airfoil for aerodynamic performance improvement

The salient aim of this paper is the shape optimization of an airfoil equipped with Gurney-Flap for aerodynamic performance improvement. The optimization of the slotted Gurney-Flap for improving the aerodynamic efficiency and increasing the lift force of NACA 0012 is the novelty of this research. The Genetic Algorithm, Artificial Neural Network, and Computational Fluid Dynamics are employed for shape optimization. The optimization variables include the height and thickness of the GF; also, the thickness and position of the slot. All analyses have been conducted at Re=0.45×106 and AOA=8o to simulate the take-off phase. After validating numerical results, the optimization process was conducted with two different fitness functions of lift force and aerodynamic efficiency. According to the results, the geometry representing an optimized lift coefficient compared to the geometry with optimized aerodynamic efficiency has a considerably higher height. Furthermore, the thickness of the slot in the first geometry is lower than the second geometry. The first optimized geometry leads to a 21.64 percent lift coefficient increment; there is a 293 percent increment in aerodynamic efficiency due to the second optimized geometry. As a result, it can be indicated that the help of slotted optimized Gurney-Flap can provide the required lift force for Short Take-Off Landing distance.

Computational evaluation of geometric effects on aerodynamic performance of circulation control airfoils

The geometric effects of Coanda trailing edges on the aerodynamic performance of an airfoil are numerically evaluated for a range of different freestream Mach numbers and momentum coefficients. A Circulation control (CC) airfoil with a circular trailing edge (ACTE) proves to have better control effectiveness at low subsonic freestream speeds (Mach = 0.1). A CC airfoil having an elliptic trailing edge (AETE) outperforms the ACTE at high subsonic flow conditions. The occurrence of C μ-stall for the AETE is greatly postponed, and meanwhile the maximum net lift coefficient increment achieved for the AETE (Δ C L = 0.51) is slightly higher than that of the ACTE (Δ C L = 0.50) at Mach 0.6. Compared to the ACTE, the AETE is found to have better control consistency at different operating velocities and better control stability when the Coanda jet is supersonic. Through careful consideration of the aerodynamic performance and the control effects, the most appropriate axial ratio for an AETE ellipse is within the interval from 1.5 to 2. Finally, the flow field instability phenomenon and the jet detachment induced by the supersonic Coanda jet are investigated. A self-sustained shock-wave instability phenomenon without jet detachment is first observed in this paper.

AERODYNAMIC STUDY OF SLOTTED FLAP FOR NACA 24012 AIRFOIL BY DYNAMIC MESH TECHNIQUES AND VISUALIZATION FLOW

Slotted flap is one of high lift devices. It considered as a moving part of the airfoil which is used as a control instrument in a form of elevator, rudders and ailerons. The main focus of work is to investigate the effect of flap chord, gap and overlap on aerodynamic characteristic of NACA 24012 airfoil. The model was tested with 20% C, 30% C and 40% C single slotted flaps at zero angle of attack. The dynamic mesh and user defined function is applied to control the flap distance with respect to wing at any position. The simulation was done by solving the governing equations (Continuity, Reynolds Averaging Naveir- Stokes and Energy Equation) in 2-D using Fluent analysis at Reynolds number of 3.1x〖10〗^6. Based on the results presented, larger increment of lift coefficient is obtained with the larger flap chord, but this increase is accompanied by a drag penalty. Furthermore, the loss of lift coefficient associated with larger extending flap at 3% C had a very detrimental effect on the attainable lift coefficient. The simulation result also shows that an optimum gap is 1% C in order to derive the maximum lift capability from the flap model. The code is validated against field measurements to show how close the CFD model simulates the reality.

Aerodynamic characterization of bio-mimicked pleated dragonfly aerofoil

The work inspired by the dragonfly wing corrugation positioned at the front wing's radius section lying at 40% of the total wingspan of forewing from the root section. During gliding flight, dragonfly wings presumed to be an ultra-light aerofoil due to its well-defined cross-sectional corrugation. The aerodynamic simulation carried out to understand the aerodynamic performance of a bio-mimetic dragonfly corrugated airfoil at low Reynolds number range of 75000-150000 to explore the potential advantages of pleated airfoils at a varying angle of attack from 0° to 12°. CFD analysis accomplished by using ANSYS Fluent to understand the aerodynamic performance of the pleated aerofoil and its counterpart profiled aerofoil. The simulation result exhibits aerodynamic commission, i.e. the corrugated aerofoil's gliding ratio (CL/CD), is higher than that of a profiled aerofoil. The analysis demonstrates that the pleated corrugated aerofoil produces a comparable higher lift coefficient than that of profiled aerofoil NACA 0015 and reduces the separation bubble. The vortices trapped inside the valleys, which re-energize the flow and delay flow separation leads to an increment in lift coefficient (CL). These investigations amplify the demand of insect-inspired corrugated wing structure and make a possible application for improved design of pretended wings for micro air vehicles.

Laminar flow instabilities of a grooved circular cylinder

This paper represents a numerical study on the flow past a grooved circular cylinder with a diameter of D. The study aims to investigate the instabilities of the flow at laminar flow regimes when Reynolds number varies from 50 to 200. Three different geometrical grooves are examined, including square, triangular, and dimple. The grooves are located at $$ \theta $$ = 0°, 30°, 45°, 60° and 90°, separately relative to the front stagnation point. The effects of the groove on the flow control and imposed forces on the cylinder are studied at different Re. It is found that the presence of the groove affects the dynamic behaviour of the streamwise vortices. This behaviour leads to alteration of the pressure coefficient and a significant increment in lift coefficient, in particular for the triangular grooved cylinder. At Re = 200, the results show that the magnitude of the lift coefficient of the triangular grooved cylinder with $$ \theta $$ = 45° is 200% more than that of the square grooved cylinder. The results also reveal that the drag coefficient reduces up to 100% by employing a triangular groove shape at $$ \theta $$ = 45°. These variation in lift and drag coefficients leads to maximising the aerodynamic efficiency factor (lift-to-drag ratio) of the triangular grooved cylinder up to 200%.

Testing of High-Lift Common Research Model with Integrated Active Flow Control

A 10%-scale high-lift version of the Common Research Model and an active flow control (AFC) version of the model equipped with a simple-hinged flap were successfully tested. The main objective of the test was to develop an AFC system that can provide the necessary lift recovery on a simple-hinged flap high-lift system while minimizing its pneumatic power requirement. The tests were performed in the 14- by 22-Foot Subsonic Tunnel at the NASA Langley Research Center. Three types of AFC devices were examined: the double-row sweeping jets (DRSWJ), the alternating pulsed jets (APJ), and the high-efficiency low-power (HELP) actuators. The DRSWJ and the APJ actuators used two rows of unsteady jets, whereas the HELP actuators used a combination of unsteady and steady jets, to overcome strong adverse pressure gradients while minimizing the mass flow usage. Nozzle pressure ratio, mass flow consumption, and the power coefficient were used for judging the performance efficiency of the AFC devices. The current data are presented with the Transonic Wall Interference Correction System method applied. The HELP actuation concept was effective in controlling flow separation in the linear region of the curves comparing lift coefficient to mass flow rate. The HELP actuation achieved a targeted lift coefficient increment of 0.5, thus meeting the lift performance goal of the NASA Advanced Air Transport Technology project.

Effects of circular trailing edge with the dimple on flow separation of NACA S1210 hydrofoil

Flow separation over a hydrofoil affects its hydrodynamic performance and ultimately reduces the power production of a turbine. Therefore, the different flow control methods have been proposed by the researchers to overcome these challenges. In this study, the combined effects of circular trailing edge with dimple have been investigated numerically on NACA S1210 hydrofoil to control the flow separation of the hydrofoil. The results showed that the coupled effects of the dimple and circular trailing edge modification appeared effective compared to baseline hydrofoil at a wide range of angles of attack. Maximum lift coefficient increments are 9 and 12% at an angle of attack of 10° and 12°, respectively, for circular trailing edge hydrofoil. The maximum increase of glide ratio for the coupled effect (circular trailing edge with outward dimple) is approximately 114% at an angle of attack of 12°. The outcome will be advantageous to design a proper blade profile for horizontal axis current turbines.

Serrations effect on the aerodynamic performance of wind turbine airfoils

A numerical and experimental analysis of the trailing edge serrations effect on the aerodynamic performance of the NACA643418 airfoil is presented. 3D Reynolds-averaged Navier-Stokes simulations (RANS) of the flow over the airfoil with and without trailing edge serrations have been carried out with the Transitions SST turbulence model. From these calculations an empirical law for the prediction of the aerodynamic behaviour of the airfoil with serrations is derived. The law is validated through wind tunnel experiments. The NACA643418 airfoil has been tested with trailing edge serrations installed with different flap angles to analyze the effect of the parameter on the performance of the airfoil. Results show that the trailing edge serrations cause a significant increment in lift coefficient with low penalties in drag.

Effect of relative camber on the aerodynamic performance improvement of asymmetrical blunt trailing-edge modification

In this paper, the aerodynamic performance of the S series of wind turbine airfoils with different relative cambers and their modifications is numerically studied to facilitate a greater understanding of the effects of relative camber on the aerodynamic performance improvement of asymmetrical blunt trailing-edge modification. The mathematical expression of the blunt trailing-edge modification profile is established using the cubic spline function, and S812, S816 and S830 airfoils are modified to be asymmetrical blunt trailing-edge airfoils with different thicknesses. The prediction capabilities of two turbulence models, the k-ω SST model and the S-A model, are assessed. It is observed that the k-ω SST model predicts the lift and drag coefficients of S812 airfoil more accurately through comparison with experimental data. The best trailing-edge thickness and thickness distribution ratio are obtained by comparing the aerodynamic performance of the modifications with different trailing-edge thicknesses and distribution ratios. It is, furthermore, investigated that the aerodynamic performance of original airfoils and their modifications with the best thickness of 2% c and distribution ratio being 0:4 so as to analyze the increments of lift and drag coefficients and lift–drag ratio. Results indicate that with the increase of relative camber, there are relatively small differences in the lift coefficient increments of airfoils whose relative cambers are less than 1.81%, and the lift coefficient increment of airfoil with the relative camber more than 1.81% obviously decreases for the angle of attack less than 6.3°. The drag coefficient increment of S830 airfoil is higher than that of S816 airfoil, and those of these two airfoils mainly decrease with the angle of attack. The average lift–drag ratio increment of S816 airfoil with the relative camber of 1.81% at different angles of attack ranging from 0.1° to 20.2° is the largest, closely followed by S812 airfoil. The lift–drag ratio increment of S830 airfoil is negative as the angle of attack exceeds 0.1°. Thus, the airfoil with medium camber is more suited to the asymmetrical blunt trailing-edge modification.

The research on active flow control method with vibration diaphragm on a NACA0012 airfoil at different stalled angles of attack

The flow control method with vibration diaphragm as a relatively new active flow control method has shown its great potential of improving flow field. A two-dimensional numerical computation for flow over an airfoil with vibration diaphragm is present at different stalled angles of attack and Re=1.2×105. The vibration diaphragm is placed at the suction side of the airfoil surface near the leading edge and the length of the vibration diaphragm's projection along the chord direction is 0.1c. The effects of the dependant parameters dimensionless amplitude and frequency and the optimum parameters varying with different stalled angles of attack are investigated through enormous simulation data. Numerical results show that the lift coefficient and the lift-to-drag ratio increase firstly with the growth of the dimensionless frequency and then decrease. There exists the optimum dimensionless amplitude to make the control method behave best. The maximum increment of lift coefficient occurs at the low dimensionless frequency near 0.5, whereas for the lift-to-drag ratio, the maximum increment is attained the value of dimensionless frequency ranging from 1 to 1.5 at 16 degree angle of attack, the greatest improvement of the lift-to-drag ratio occurs when the value of dimensionless amplitude and frequency are 0.015 and 1.4, the maximum percent improvement is 91.31%. The oscillation coefficient is found to be a critical parameter to affect the lift-to-drag ratio of the airfoil and there exits an optimum oscillation coefficient to obtain greatest improvement of lift-to-drag ratio. The optimum parameters of the control methods to achieve the highest lift-to-drag ratio vary with different stalled angles of attack, with the growth of the angles of attack, the optimum dimensionless amplitude increases, the optimum dimensionless frequency decreases firstly and then keep the certain value 1.1 and the oscillation coefficient fluctuates weakly in the range of 1.15–1.3.

Separation control using quasi-radial wall jets

In this paper, flow separation control over a NACA 4415 wing section using arrays of discrete quasi-radial wall jets is numerically investigated. This novel flow control method is studied under conditions which the wing section angle of attack is α=18 degrees and Reynolds number based on chord length is Rec=550000. Contours of mean streamwise velocity indicated that quasi-radial wall jets cover wing surface rapidly. Therefore, arranging them in a spanwise row can be an effective way in adding momentum to the near wall fluid particles. According to the results, if blowing parameters of the jets be chosen correctly, arrays of quasi-radial wall jets would lead to considerable decrease in wake region extent and would bring enhancement of wing aerodynamic performance (up to 46% increment of lift coefficient). Moreover, investigation of streamlines and velocity vectors reveals the presence of two helical vortices on both edges of quasi-radial wall jets. These vortices, which form due to interactions between external stream and spanwise velocity component of jets, generate mixing in the surrounding fluid. It could be said that besides momentum addition, quasi-radial wall jets in an external stream act like vortex generators.

Wind tunnel experiments on flow separation control of an Unmanned Air Vehicle by nanosecond discharge plasma aerodynamic actuation**Project supported by the National Natural Science Foundation of China (Grant Nos. 61503302, 51207169, and 51276197), the China Postdoctoral Science Foundation (Grant No. 2014M562446), and the Natural Science Foundation of Shaanxi Province, China (Grant No. 2015JM1001).

Plasma flow control (PFC) is a new kind of active flow control technology, which can improve the aerodynamic performances of aircrafts remarkably. The flow separation control of an unmanned air vehicle (UAV) by nanosecond discharge plasma aerodynamic actuation (NDPAA) is investigated experimentally in this paper. Experimental results show that the applied voltages for both the nanosecond discharge and the millisecond discharge are nearly the same, but the current for nanosecond discharge (30 A) is much bigger than that for millisecond discharge (0.1 A). The flow field induced by the NDPAA is similar to a shock wave upward, and has a maximal velocity of less than 0.5 m/s. Fast heating effect for nanosecond discharge induces shock waves in the quiescent air. The lasting time of the shock waves is about 80 μs and its spread velocity is nearly 380 m/s. By using the NDPAA, the flow separation on the suction side of the UAV can be totally suppressed and the critical stall angle of attack increases from 20° to 27° with a maximal lift coefficient increment of 11.24%. The flow separation can be suppressed when the discharge voltage is larger than the threshold value, and the optimum operation frequency for the NDPAA is the one which makes the Strouhal number equal one. The NDPAA is more effective than the millisecond discharge plasma aerodynamic actuation (MDPAA) in boundary layer flow control. The main mechanism for nanosecond discharge is shock effect. Shock effect is more effective in flow control than momentum effect in high speed flow control.

Numerical simulation of the effect of relative thickness on aerodynamic performance improvement of asymmetrical blunt trailing-edge modification

In this paper, the aerodynamic performance of wind turbine airfoils with different relative thicknesses and their modifications has been numerically investigated to facilitate a greater understanding of the effects of maximum relative thickness and its position on the aerodynamic performance improvement of asymmetrical blunt trailing-edge modification. The lift and drag coefficients of airfoil NACA4415 are calculated with the k-ω SST turbulence model, and are compared with experimental data to validate the simulation accuracy of the Computational Fluid Dynamics (CFD) approach. The airfoils with different relative thicknesses are modified to be asymmetrical blunt trailing-edge airfoils by means of the software Xfoil. The best trailing-edge thickness distribution ratio is obtained by comparing the aerodynamic performance of the modifications with different distribution ratios. The aerodynamic performance of original airfoils and their asymmetrical modifications with the best thickness distribution ratio being 1:3 is investigated to analyze the increments of lift and drag coefficients and lift-drag ratio. Results indicate that with the increasing of relative thickness, the lift coefficient increment of NACA4418 airfoil is the smallest for the angle of attack more than 9°, and the drag coefficient increment as a whole decreases first and then increases, but the average lift-drag ratio increment of NACA4412 airfoil is the largest, closely followed by NACA4415 airfoil. It is also showed that with the relative thickness position close to the leading-edge, the increments of lift and drag coefficients decrease and increase for the angle of attack more than a certain value, respectively, and the average lift-drag ratio increment of NACA4415 airfoil is positive and larger than those of NACA4415-mod25 and NACA4415-mod20 airfoils. Therefore, the medium thickness airfoil whose relative thickness position is away from the leading-edge is more suited to the asymmetrical blunt trailing-edge modification.

Experimental Aerodynamic Characteristics of a Compound Wing in Ground Effect

Wing configuration is a parameter that affects the performance of wing-in-ground effect (WIG) craft. In this study, the aerodynamic characteristics of a new compound wing were investigated during ground effect. The compound wing was divided into three parts with a rectangular wing in the middle and two reverse taper wings with anhedral angle at the sides. The sectional profile of the wing model is NACA6409. The experiments on the compound wing and the rectangular wing were carried to examine different ground clearances, angles of attack, and Reynolds numbers. The aerodynamic coefficients of the compound wing were compared with those of the rectangular wing, which had an acceptable increase in its lift coefficient at small ground clearances, and its drag coefficient decreased compared to rectangular wing at a wide range of ground clearances, angles of attack, and Reynolds numbers. Furthermore, the lift to drag ratio of the compound wing improved considerably at small ground clearances. However, this improvement decreased at higher ground clearance. The drag polar of the compound wing showed the increment of lift coefficient versus drag coefficient was higher especially at small ground clearances. The Reynolds number had a gradual effect on lift and drag coefficients and also lift to drag of both wings. Generally, the nose down pitching moment of the compound wing was found smaller, but it was greater at high angle of attack and Reynolds number for all ground clearance. The center of pressure was closer to the leading edge of the wing in contrast to the rectangular wing. However, the center of pressure of the compound wing was later to the leading edge at high ground clearance, angle of attack, and Reynolds number.

Optimization of Active/Passive Flow Control Parameters on Airfoils at Transonic Speeds

To improve aerodynamic performance, the active and passive flow control parameters are optimized in different combinations. The flow is simulated by solving the Navier–Stokes equations and a turbulence model. The active flow control is applied using flow suction or blowing on an airfoil, and its design variables are set to be the location and the angle of the suction/blowing port and the mass flow coefficient at this port. Airfoil surfaces are parameterized by using Bezier curves to optimize an airfoil shape in a passive way. To validate the computational methodology, successful comparisons have beenmadewith available data. A reel coded genetic algorithm is used in all optimization processes. For the demonstration problems considered herein, the best result is provided by the sequential optimization process, which results in 575.8% increment in aerodynamic performance, 183.76% increment in lift coefficient, and 68% decrease in drag coefficient.

Experimental Study on Gurney Flap and Apex Flap on Delta Wing

The effects of Gurney flaps and apex flaps on the longitudinal aerodynamic performance of a delta wing with 70-deg sweepback angle were investigated in low-speed wind tunnel at a Reynolds number of 3.16 × × 10 5 based on the root chord of the delta wing. The experimental results indicate that the increment of lift coefficient can be obtained efficiently for different ranges of angle of attack with Gurney flaps and apex flaps, respectively. The study first tested the application of Gurney flaps together with an apex flap on the highly swept-back delta wing. The results show that they, used together, can greatly improve the longitudinal aerodynamic performance of the delta wing, and in this case the apex flap contributes much to the gains. The maximum lift-to-drag ratio can be achieved when the drooping angle of the apex flap is nearly equal to the angle of attack.

Data Item predicts wing lift coefficient increment at zero angle of attack due to the deployment of trailing-edge split flaps at low speeds

Data Item predicts wing lift coefficient increment at zero angle of attack due to the deployment of trailing-edge split flaps at low speeds

Analytical determination of propeller performance degradation due toice accretion

A computer code has been developed which is capable of computing propeller performance for clean, glaze, or rime iced propeller configurations, thereby providing a mechanism for determining the degree of performance degradation which results from a given icing encounter. The inviscid, incompressible flow field at each specified propeller radial location is first computed using the Theodorsen transformation method of conformal mapping. A droplet trajectory computation then calculates droplet impingement points and airfoil collection efficiency for each radial location, at which point several user-selectable empirical correlations are available for determining the aerodynamic penalities which arise due to the ice accretion. Propeller performance is finally computed using strip analysis for either the clean or iced propeller. In the iced mode, the differential thrust and torque coefficient equations are modified by the drag and lift coefficient increments due to ice to obtain the appropriate iced values. Comparison with available experimental propeller icing data shows good agreement in several cases. The code's capability to properly predict iced thrust coefficient, power coefficient, and propeller efficiency is shown to be dependent on the choice of empirical correlation employed as well as proper specification of radial icing extent.