Maximum Lift Coefficient Research Articles

Analysis on aerodynamic characteristics of morphing wing with flexible trailing edge

The wing applying active variable camber trailing edge technology can improve aircraft's aerodynamic performance,and the study of aerodynamic characteristics is of great significance to the design of the morphing wing. Choosing the 2D variable camber wing with flexible trailing edge as the object of study,the compressible Navier-Stokes equation and Spalart-Allmaras( S-A) turbulence model were applied to investigate the aerodynamics of morphing wing numerically by using the software Fluent. Aerodynamic characteristics were studied from the perspective of pressure distribution,flow field structure and the deformation mode. The numerical results show that the morphing wing has almost the same slope of lift curve and maximum lift coefficient with the conventional wing with simple flap,but has lower stall angle of attack. It has higher lift coefficient,lift-to-drag ratio,and higher nose-down pitching moment coefficient than conventional wing before the occurrence of stall. When flexible trailing edge deflects downward to a special angle,trailing edge vortex is suppressed to expand forward. The lift coefficient slowly rises after the stall,which increases the range of effective angle of attack of morphing wing and leads to a good stall characteristic.

Comparison of wing–propeller interaction in tractor and pusher configuration

This research compared a tractor and a pusher configuration for a tilt-body vertical take-off and landing micro air vehicle using an experiment in a subsonic wind tunnel. All tests were conducted in the range 6–10 m/s and 6000–8000 RPM of the freestream and the propeller rotation, respectively. The wing model was a rectangular platform with an NACA 0012 airfoil and AR = 1. The incidence angles varied within the range 0°–90° to cover flight configuration from cruise to hover and vice versa. Basically, the downstream of the propeller is stronger than the upstream. Thus the wing submerses in the propeller’s downstream, which is the tractor configuration, and has an advantage over the other configuration by increasing the freestream velocity and decreasing the angle of attack. The results of the experiment found that the wing aerodynamics of the tractor were improved by the prop-wash effect with increases of about 1–1.13 times in the wing lift curve slope and about 0.9–1.12 times in the factor K of the drag polar curve. However, in the case of the pusher, the change was not significant. The stall angle, the maximum lift coefficient, and the drag coefficient at zero angle of the wing all increased in both the tractor and the pusher configurations due to the prop-wash effect. Moreover, after experiencing the prop-wash effect, the aerodynamic center of the tractor wing was almost at the same point.

Experimental investigations for parametric effects of dual synthetic jets on delaying stall of a thick airfoil

A promising strategy of synthetic jet arrays (SJA) control for NACA0021 airfoil in preventing flow separation and delaying stall is investigated. Through aerodynamic forces, flowfield and velocity profiles measurements, it indicates that the synthetic jet (SJ) could enlarge the mixing of the shear layer and then enhance the stability of boundary layer, resulting in scope reduction of the flow separation zone. Furthermore, the control effects of dual jet arrays positioned at 15%c (Actuator 1) and 40%c (Actuator 2) respectively are systematically investigated with different jet parameters, such as two typical relative phase angles and various incline angles of the jet. The jet closer to the leading edge of airfoil is more advantageous in delaying the stall of airfoil, and overall, the flow control performances of jet arrays are better than those of single actuator. At the angle of attack (AoA) just approaching and larger than the stall AoA, jet array with 180° phase difference could increase the lift coefficient more significantly and prevent flow separation. When momentum coefficient of the jet arrays is small, a larger jet angle of Actuator 2 is more effective in improving the maximum lift coefficient of airfoil. With a larger momentum coefficient of jet array, a smaller jet angle of Actuator 2 is more effective.

Impact of Gurney Flaplike Strips on the Aerodynamic and Vortex Flow Characteristic of a Reverse Delta Wing

The impact of Gurney flaplike strips, of different geometric configurations and heights, on the aerodynamic characteristics and the tip vortices generated by a reverse delta wing (RDW) was investigated via force-balance measurement and particle image velocimetry (PIV). The addition of side-edge strips (SESs) caused a leftward shift of the lift curve, resembling a conventional trailing-edge flap. The large lift increment overwhelmed the corresponding drag increase, thereby leading to an improved lift-to-drag ratio compared to the baseline wing. The lift and drag coefficients were also found to increase with the strip height. The SES-equipped wing also produced a strengthened vortex compared to its baseline wing counterpart. The leading-edge strips (LESs) were, however, found to persistently produce a greatly diffused vortex flow as well as a small-than-baseline-wing lift in the prestall α regime. The downward LES delivered a delayed stall and an increased maximum lift coefficient compared to the baseline wing. The LESs provide a potential wingtip vortex control alternative, while the SESs can enhance the aerodynamic performance of the RDW.

Aerodynamic Characteristics of Low-Aspect-Ratio Wings with Various Aspect Ratios in Low Reynolds Number Flows

The effect of the aspect ratio on the stall characteristics of low-aspect-ratio wings is experimentally investigated at a Reynolds number of 5.2×104. The aspect ratio ranges from 0.5 to 1.5 at intervals of 0.1. The aerodynamic coefficients of thin rectangular wings are measured in a low-speed wind tunnel. It is found that the aerodynamic characteristics of low-aspect-ratio wings are significantly sensitive to the aspect ratio. The maximum lift coefficient increases as the aspect ratio decreases toward unity. On the other hand, the difference in lift-to-drag ratio is negligible at large angles. Hysteresis during a stall occurs when the aspect ratio is within a narrow range. The span of the hysteresis loop has its peak at an aspect ratio of 1.0. The discussion also includes the analysis of lift components and flow visualization.

Global Stability Analysis of a Compressible Turbulent Flow Around a High-Lift Configuration

The stability of flow over a complex high-lift configuration with significant regions of separated flow is analyzed. Current state-of-the-art flow solvers encounter difficulties in predicting both the onset of flow separation over similar configurations and the progression of the separated region when the angle of attack is increased. The present analysis provides useful insights into the mechanisms responsible for limiting the maximum lift on multielement wing configurations by relating flow separation with the onset of flow instability. A steady solution based on a finite volume discretization is used as the basic state for the linear stability analysis. The resulting generalized eigenvalue problem has been solved using a Krylov subspace projection technique in the form of the Arnoldi iterative method. This methodology is first applied to a NACA0012 test case at subsonic and transonic conditions. Then, for the first time, the stability of flow over an industrial multicomponent geometry involving the A310 airfoil has been investigated in order to identify the low-frequency instabilities related to flow separation. One important conclusion is that, for both configurations, one unstable mode related to flow separation appears at an angle of attack greater than that corresponding to the maximum lift coefficient condition. Finally, the adjoint problem is solved in order to evaluate the receptivity characteristics and the structural sensitivity of both configurations, identifying the regions of flow that can be manipulated to yield the largest change in the flowfield.

Experimental investigation of aerodynamic performance of oscillating wings at low Re numbers

In this study, the effect of varying aspect ratios on aerodynamic performances on the flat plate and NACA 2415 airfoil without/with low frequency pitching movement at low Reynolds (Re) numbers was investigated. The experiments include force measurements for both static and dynamic airfoils and smoke-wire flow visualization for static airfoils at Re numbers of 50,000, 75,000 and 100,000. In the static cases, the stall delayed because of tip vortices at low aspect ratios, as the aspect ratio increased, maximum lift coefficient and the stall angle decreased. On the other hand, the stall mechanism changed and the abrupt stall occurred. As the Re number increased, the mild stall formed. In the dynamic cases, on the flat plates at lower ARs, the oscillation did not change the lift coefficients considerably, and as the Re number increased, the oscillation made the drag force decrease. In the dynamic NACA 2415 wing with AR1, the oscillation improved the aerodynamic performance at down-stroke conditions (especially between 15° and 25°). However, in dynamic NACA 2415 wing with AR2, as the Re number increased, dynamic stall formed over the wings at the angle of attacks between 15° and 25°.

A new direct design method of wind turbine airfoils and wind tunnel experiment

A general parametric representation function (Shape Function) for wind turbine – dedicated airfoils based on Taylor high-order polynomial series is presented for the first time. The design space and shape control function of the airfoil have been studied. The objective of the high-performance WT (Wind Turbine) airfoils was to maximise the lift/drag ratio at the design angle of attack both in free and fixed transitions. The optimised mathematical model of the airfoils is built combing genetic algorithm and the flow solver RFOIL. The new airfoil family (ranging in thickness from 15 to 20%) with target characteristics were designed for variable-speed operation with pitch control of large megawatt-sized rotors. Wind tunnel experiments for the WT180 airfoil were carried out for both clean and rough conditions. The experimental results for the lift and drag coefficients agree well with the RFOIL predictions. The testing verified the high lift/drag ratio and the high maximum lift coefficient for the WT180 airfoil. The results indicate that this novel design method is feasible to optimise wind turbine airfoils.

Repetitively enhanced neural networks method for complex engineering design optimisation problems

Abstract A Repetitively Enhanced Neural Networks (RENN) method is developed and presented for complex and implicit engineering design problems. The enhanced neural networks module constructs an accurate surrogate model and avoids over-fitting during neural networks training from supervised learning data. The optimiser is executed by the enhanced neural networks models to seek a tentative optimum point. It is repetitively added into the supervised learning data set to refine the surfaces until the RENN tolerance is reached. The RENN method demonstrates the effectiveness and feasibility for a 2D highly non-linear numerical example and the structure design of a two-member frame reaching a convergent solution at 10 and 15 iterations at the maximum error of 1% when compared with the exact solution. Then, the RENN method is applied for a long endurance unmanned aerial vehicle (UAV) aerofoil design optimisation. A Class/Shape function transformation (CST) geometry parameterisation method represents an accurate UAV aerofoil with ten geometry design variables. The high-fidelity analysis solver with structured mesh is used for a UAV aerofoil design problem. Using the RENN method, an optimal UAV aerofoil is obtained using 88 high fidelity evaluations at an error of 1·24%. The process reduces the computational time by 81·2% compared with the full high fidelity model. The optimal aerofoil shows a drag reduction of 10·8% in the cruise condition and an improvement in the maximum lift coefficient and stall angle-of-attack when compared with the baseline AG24 aerofoil.

A Simple and Robust Subsonic Boundary Layer Method to Be Used with the Panel Method for Wing Calculations Using a Wing Strip Approach

This paper will present the development of a simple subsonic boundary layer method suitable to be used coupled with panel methods in order to estimate the aerodynamic characteristics, including viscous drag and maximum lift coefficient, of 3D wings. The proposed method does not require viscous-inviscid iterations and is based on classical integral bi-dimensional boundary layer theory using Thwaites and Head ́s models with bi-dimensional empirical corrections applied to each wing strip being therefor robust and efficient to be used in the early conceptual stage of aircraft design. Presented results are compared to the Modified CS Method in an IBL scheme and experimental data and are shown to provide good results.

A New Design Concept for Wind Turbine Airfoil

Wind energy has been attracting more and more attentions due to its clean and renewable source. The aerodynamic characteristic of wind turbine airfoil directly affects the turbine efficiency. In order to improve the airfoil aerodynamic characteristic, a new concept airfoil configuration for wind turbine is presented. A cave on the upper surface near the trailing edge is designed to generate a trapped vortex in the cave. The trapped vortex is used to stabilize the separated flow when the airfoil at high angle of attack. Combining with the Gurney flap, the airfoil with the cave behaves very good aerodynamic characteristics at wide range of incidences, especially at high angles of attack. The method is used on the well-known FFA-W3-301 turbine airfoil. By using numerical simulation, it is shown that the new airfoil has a higher lift than the original airfoil at the same angle of attack, the stall angle of attack increases from 12 degree to 17 degree, and the maximum lift coefficient increases approximately 64 percents. In addition, the effects of the chord-wise location of starting point of the designed cave are discussed. Therefore, it is believed that the new-designed concept can be investigated and explored further for wind turbine.

Vortex Shedding and Aerodynamic Performance of Airfoil with Multiscale Trailing-Edge Modifications

An experimental investigation was conducted into the nature of the vortex shedding generated by truncated and nonflat serrated trailing edges of a NACA 0012 wing section as well as the aerodynamic performance of such trailing edges. In line with previous findings, the truncated trailing edge generates a significant amount of vortex shedding, while increasing both the maximum lift and drag coefficients, resulting in an overall reduction in the maximum lift-to-drag ratio compared to a plain NACA 0012 wing section. By decreasing the chevron angle of the nonflat trailing-edge serrations (i.e., by making them sharper), the energy of the vortex shedding significantly decreases, and the lift-to-drag ratios increase compared to a plain NACA 0012 wing section. Fractal/multiscale patterns made of scaled-down repetitions of these serrations were also investigated with a view to further improve performance. It was found that the energy of the vortex shedding increases with an increasing fractal iteration if the chevron is broad () but decreases for sharper chevrons (). It is believed that if is too big, the multiscale trailing edges are too far away from each other to interact and break down the vortex shedding mechanism. Fractal/multiscale trailing edges are also able to improve aerodynamic performance compared to the NACA 0012 wing section.

The numerical simulation of the wing kinematics effects on near wake topology and aerodynamic performance in hovering Drosophila flight

The parallel large-scale unstructured finite volume method based on an Arbitrary Lagrangian–Eulerian (ALE) formulation in Erzincanli and Sahin (2013) has been employed in order to investigate the effects of the wing kinematics on the near wake topology produced by a pair of flapping Drosophila wings in hover flight. The three-dimensional wing shape and the baseline wing kinematics are the same as those of the robotic fly by Dickinson et al. (1999). The simulations are used to quantify the important wing kinematic parameters for the wake topology and as well as their correlations with the force production. The angle-of-attack is shown to be very effective for producing lift during the wing translational motion. However, for larger values the angle-of-attack limits the angle during the stroke reversal and reduces the rotational lift. The angle-of-attack at which the maximum lift coefficient occurs is in excellent agreement with the earlier experiments in the literature. In addition, the numerical results confirm that the increase in the wing stroke amplitude leads to a prolonged attachment of the leading edge vortex (LEV) over a relatively large distance and increases force production. The variations in the heave angle include the figure-of-eight pattern, figure-of-O pattern and figure-of-U pattern. The constant heave angle and figure-of-eight pattern are found to have a more profound influence on the magnitude of force production. In addition, some asymmetric variations in the wing kinematics are introduced in order to assess the important parameters determining forward and backward flights. The paddling wing motion is shown to be very effective to initiate forward and backward acceleration.

Effects of propeller thrust trimming on tailless unmanned aerial vehicle performance

Flight performance and longitudinal static stability are two contradictory requirements for aircraft. This is especially true for a tailless configuration, whose flight performance is traded for static stability. A new tailless gull wing configuration with propeller thrust trimming is proposed in this paper. It is expected that the propeller thrust can help to provide a nose-up pitching moment and improve the flight performance. In the present configuration, a cambered airfoil is used to increase the lift-to-drag ratio as well as the maximum lift coefficient. Two approaches are adopted to obtain the nose-up pitching moment. Firstly, the thrust line is shifted below the center of gravity to make a positive contribution to the pitching moment. Secondly, the thrust vectoring is employed to trim longitudinally instead of elevon deflection. The wind-tunnel data indicate that the lift-to-drag ratio and maximum lift coefficient increase notably with the help of propeller thrust trimming. To determine the performance benefits of the gull wing with propeller thrust trimming, a conceptual design environment of the electric-powered unmanned aerial vehicle is developed by integrating the modules of the conceptual parameters, aerodynamics, stability, propulsion, weight, performance, as well as optimization. The optimization results suggest that, compared with the two other conventional tailless configurations, the gull wing with propeller thrust trimming results in 20 and 34% endurance increase. Flight tests of a 2.5 kg built aircraft demonstrate the aerodynamic feasibility of the gull wing with propeller thrust trimming concept.

LIFT PERFORMANCE OF A TWIST MORPHING MAV WING

The manoeuvrability performance of a twist morphing MAV has been the main interest for the past researches. However, aerodynamic behaviour of a twist morphing wing is not fully explored due to limited MAV wing size, limited energy budgets, complicated morphing mechanism, and complex aerodynamic-wing structural interaction. Therefore, the effect of a twist morphing wing mobility on the lift distribution of MAV wing is still remained unknown. Thus, present work was carried out to compare the lift performance between a twist morphing wing with membrane and rigid MAV wing design. A quasi-static aeroelastic analysis by using the Ansys-Fluid Structure Interaction (FSI) method is utilized in current works to calculate the lift performance for each MAV wing design. Each MAV wing has identical wing dimension except for twist morphing wing where a 3N morphing force was imposed on the wing to produce the twist mobility. The lift results show that twist morphing wing able to produce (5% to 20%) higher lift magnitude compared to the membrane and rigid wing for every angle attack cases at pre-stall angle. However, twist morphing wing had slightly suffered from (at least 1°) earlier stall angle and produced almost similar maximum lift coefficient magnitude to the membrane wing

Lift and Drag Characteristics of an Air-Cooled Heat Exchanger Axial Flow Fan

Actuator-disk models (ADMs) use blade element theory to numerically simulate the flow field induced by axial fans. These models give a fair approximation at near design flow rates, but are of poor accuracy at low flow rates. Therefore, the lift/drag (LD) characteristics of two-dimensional (2D) sections along the span of an air-cooled heat exchanger (ACHE) axial fan are numerically investigated, with the future prospect of improving ADMs at these flow conditions. It is found that the blade sectional LD characteristics are similar in shape, but offset from the 2D LD characteristics of the reference airfoil (NASA LS 413 profile) at small angles of attack (αatt<5deg). A deviation between these characteristics is observed at higher angles of attack. The blade sectional lift coefficients for αatt>5deg always remain lower compared to the maximum lift coefficient of the reference airfoil. Conversely, the blade sectional drag coefficients are always higher compared to that of the reference airfoil for αatt>5deg.

Optimisation of a novel trailing edge concept for a high lift device

This study aimed to observe the effect of a novel concept (referred to as the flap extension) implemented on the leading edge of the flap of a three element high lift device. The high lift device, consisting of a flap, main element and slat is designed around an Airbus research profile for sufficient take off and landing performance of a large commercial aircraft. The concept is realised on the profile and numerically optimised to achieve an optimum geometry. Two different optimisation approaches based on Genetic Algorithm optimisations are used: a zero order approach which makes simplifying assumptions to achieve an optimised solution: as well as a direct approach which employs an optimisation in ANSYS DesignXplorer using RANS calculations. Both methods converge to different optimised solutions due to simplifying assumptions. The solution to the zero order optimisation showed a decreased stall angle and decreased maximum lift coefficient against angle of attack due to early stall onset at the flap. The DesignXplorer optimised solution matched that of the baseline solution very closely. The concept was seen to increase lift locally at the flap for both optimisation methods.

Investigation of flow field of clap and fling motion using immersed boundary coupled lattice Boltzmann method

This paper deals with the investigation of flow field due to clap and fling mechanism using immersed boundary coupled with lattice Boltzmann method. The lattice Boltzmann method (LBM), an alternative to Navier–Stokes solver, is used because of its simplicity and computational efficiency in solving complex moving boundary problems. Benchmark problems are simulated to validate the code, which is then used for simulating flow over two elliptic wing of aspect ratio 5 performing clap and fling flapping motion for different flow parameters such as Reynolds number (Re=75, 100, 150), advance ratio (J=10E−3.10E−2,0.2) and frequency (f=0.05Hz, 0.25Hz). Numerical simulation is able to capture typical low Reynolds number unsteady phenomena such as, ׳wake vortex wing interaction׳, ׳Kramer effect׳ and ׳delayed stall׳. The results are both qualitatively and quantitatively consistent with experimental observation. The parametric study involving different combinations of Re, f and J depict distinctly different aerodynamic performances providing physical insights into the flow physics. It is observed that a combination of low f, low J and high Re flow results in better aerodynamic performance. Pronounced lift enhancement via leading edge vortices are obtained in unsteady regime (J<1) compared to quasi-steady regime (J>1). The role of leading edge vortices in enhancing lift are investigated by studying the size and strength of these vortices for different flow conditions. For a given Re, the magnitude of maximum lift coefficient decreases with increasing f irrespective of the value of J; while the same is enhanced with the increasing Re.

A rapidly settled closed-loop control for airfoil aerodynamics based on plasma actuation

This paper presents an experimental investigation on the response of the slope seeking with extended Kalman filter (EKF) deployed in a closed-loop system for airfoil aerodynamics control. A novel dielectric barrier discharge (DBD) plasma actuator was used to manipulate the flow around the NACA 0015 airfoil. Experiments were performed under different freestream velocities U ∞, covering the chord Reynolds number Re from 4.4 × 104 to 7.7 × 104. Firstly, the advantages of applying this DBD plasma actuator (hereafter called sawtooth plasma actuator) on the airfoil were examined in an open-loop system at Re = 7.7 × 104. The sawtooth plasma actuator led to a delay in the stall angle α stall by 5° and an increase in the maximum lift coefficient $$C_{{{\text{L}}_{ \text{max} } }}$$ by about 9 %. On the other hand, at the same input power, the traditional DBD plasma actuator managed a delay in α stall by only 3° and an increase in $$C_{{{\text{L}}_{ \text{max} } }}$$ by about 3 %. Secondly, the convergence time t c of the lift force F L at Re from 4.4 × 104 to 7.7 × 104 was investigated for two closed-loop systems. It has been demonstrated that the t c was about 70 % less under the slope seeking with EKF than that under the conventional slope seeking with high-pass (HP) and low-pass (LP) filters at Re = 7.7 × 104. The reduction in t c was also observed at a different Re. Finally, the slope seeking with EKF showed excellent robustness over a moderate Re range; that is, the voltage amplitude determined by the control algorithm promptly responded to a change in Re, much faster than that of the conventional slope seeking with HP and LP filters.

Aerodynamic control of NACA 0021 airfoil model with spark discharge plasma synthetic jets

Spark discharge plasma synthetic jets (SPJs) have been used for the active flow control study on an NACA 0021 straight-wing model in a wind tunnel. The model forces and moments were measured using a six-component sting balance at a 20 m/s wind speed. The aim was to explore the SPJ’s effect on airfoil aerodynamic by examining SPJ generators’ position along the chordwise and the jet flow direction about the chord. Near the wing leading edge, two SPJ generators raised the stall angle by 2° and increased the maximum lift coefficient by 9%. The drag coefficient was decreased by 33.1%, and the lift-drag ratio was increased by 104.2% at an angle of attack above 16°. The rolling-moment coefficient was modified by 0.002, and the yawing-moment coefficient was changed by 0.0007 at angles of attack in the range of 0°–16°. The results showed that SPJs can control wing aerodynamic forces at a high angle of attack and moments at a low angle of attack.