Experimental and CFD study of slotted Krueger flaps aerodynamics in critical locations

<abstract> Ice accretion can reduce the performance of aircraft's wings, which results in higher fuel consumption and risk of accidents. Experiments proved that even in its very earlier stages (increased roughness), icing could cause a reduction of 25% in the maximum lift, and an increase of 90% in drag of an aspect ratio 6 wing. In this work, we propose a correlation to predict the degradation of the maximum lift coefficient caused by roughness effects on flows over two airfoils, a NACA 0012 and a model 5–6. In addition, a second correlation is proposed to find the minimum Reynolds number that are useful for higher Reynolds number applications when roughness is considered. The SA roughness extension is implemented into an open-source code called SU2. The verification and validation of the implementation is performed in two steps. First, the behavior of the flow over a smooth NACA 0012 is investigated to confirm whether the implementation has no influence on the original model when roughness is not activated. Then, roughness is activated, and estimations of lift coefficients and velocity profiles inside the boundary layer are evaluated and compared to numerical and experimental results. Finally, investigations on the maximum lift coefficient reduction caused by different equivalent sand grain roughness heights and Reynolds numbers are performed. Our results demonstrated that, for the equivalent sand grain roughness heights investigated, the variation of sufficiently small heights has no significant influence on the maximum lift coefficient degradation. Moreover, when roughness is continuously increased, a saturation point seems to be approached, in which the variation of the maximum lift coefficient degradation is reduced. We noticed that although the reduction of the maximum lift coefficient caused by different equivalent sand-grain roughness heights and Reynolds number present similar behavior, they fall into different curve formats. </abstract>

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.

&lt;p&gt;Flow control has significant technological importance as it can manipulate the flow field in a desired way either actively or passively. This wide research area has remained the point of attention for many years as it is applicable to various applications. Blowing as a flow control method, among other methods, is more technically feasible and market ready technique. A brief review from the existing literature on various studies on blowing has been presented along with their outcome. Then, studies were conducted to investigate the performance variation (in terms of lift coefficient, drag coefficient) of different airfoils with respect to various blowing parameters. It was observed that for NACA 0012 airfoil the maximum lift coefficient peaks at a blowing ratio 0.2 and then it decreases whereas the stalling angle increases with rise of blowing ratio. For LA203A both maximum lift coefficient and stalling angle increases with blowing ratio. For thick airfoil, it was found that lift increases with the rise of moment coefficient and blowing ratio and mid chord slots gives better performance at lower angles of attack whereas leading edge slots exhibits better performance at higher angles of attack. In case of thick elliptical airfoil, increase in lift coefficient was noticed with the increase in moment coefficient and blowing ratio but an optimum jet width to chord ratio (0.41) was found beyond which increase in jet width causes drop in lift coefficient. Additionally, it was found that at lower blowing angle thick elliptical airfoil performs better compared to that of higher angles of attack. Study on NACA 0012 and Aerospatiale A proved the fact that lifts increases with rise in blowing ratio at three different jet diameters to chord ratios. This study also found that higher blowing ratio is also advantageous for the turbo machinery in increasing the pressure difference. Lastly, the study on low Re airfoil flow found that increasing blowing ratio has a deleterious effect on aerodynamics performance (in terms of lift and drag coefficient).&lt;/p&gt;

LTHOUGH biplanes were quite popular during the early days of aviation, they had virtually disappeared from service by the mid-1930's. The biplane seemed to be plagued by inherently high drag and low maximum lift coefficient. The purpose of the present study was to re-examine the reasons for the biplane's decline. It was found that for any practical biplane configuration, it is possible to design a wing system which has a much lower weight per unit area and which has essentially the same maximum useable lift coefficient as that of a comparable, well-designed, monoplane wing, in either clean or flapped configurations. Improvements in the design of fairings permit large drag reductions, relative to the earlier designs. Because of these improvements, it is possible to design a biplane whose performance is superior, for some applications, to that of a well-designed monoplane. These applications are those for which excellent low-speed maneuverability, good short-field performance, good loadcarrying ability, low cost, and rugged construction are of primary importance.! Content If a biplane wing system is to develop a high CLmax, both wings must stall nearly together. A method presented by Fuchs was used to identify those configurations for which a good stall match could be obtained. Each wing is idealized in this method as a single horseshoe vortex, with the bound part of each vortex located at the center of pressure. Satisfactory stall match was defined as a ratio of leading wing to trailing wing Coequal to unity at an assumed CLmax for the combination. Some results of this analysis are given in Fig. 1, and some nomenclature is defined in the sketch. The airfoil chord lines are parallel; results of this study showed no advantage to be gained from the use of decalage. The geometric stagger angle a for best stall match is quite insensitive to the gap/chord ratio, and it decreases as the aspect ratio increases. The best stall match always corresponds to small values of 0, the aerodynamic stagger angle. The chordwise variation in vertical velocity induced by one airfoil upon the other corresponds to a local curvature of the flow. Since within the framework of thin airfoil theory, this effect is identical to the effect of a change in the mean camber line, it will be referred to as induced camber. Data for NACA airfoils of the four- and five-digit series show increases in the maximum section lift coefficient Cfmax

An LS(1)-0421 MOD airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3{times}5 subsonic wind tunnel (3{times}5) under steady flow and stationary model conditions, and also with the model undergoing pitch oscillations. In order to study the possible extent of performance loss due to surface roughness, a leading edge grit roughness (LEGR) pattern was developed to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers for steady state conditions were 0.75, 1, and 1.25 million, while the angle of attack ranged from {minus}10{degrees} to +40{degrees}. With the model undergoing pitch oscillations, data was acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.5 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions were used; {plus_minus} 5.5{degrees} and {plus_minus} 10{degrees}, at mean angles of attack of 8{degrees}, 14{degrees}, and 20{degrees}. For this report, unsteady conditions refer to the model in pitch oscillation. In general, the maximum unsteady lift coefficient was from 10% to 50% higher than the steady state maximum lift coefficient. Variation in the quarter chord pitching moment coefficient was nearly two times greater than steady state values at high angles of attack. These findings indicate the importance of considering the unsteady flow behavior occurring in wind turbine operation for accurate load estimates.

This paper presents the design and experimental verification of the Riso-B1 airfoil family for MW-s ize wind turbines with variable speed and pitch control . Seven airfoils were designed with thickness-to-chor d ratios between 15% and 53% to cover the entire span of a wind turbine blade. The airfoils were designed to have high maximum lift coefficient to allow a slender flexible blade while maintaining high aerodynamic efficiency. The design was carried out with a Riso inhouse multi disciplinary optimization tool. Wind tu nnel testing was done for Riso-B1-18 and Riso-B1-24 in t he VELUX wind tunnel, Denmark, at a Reynolds number of 1.6 ×10 6 . For both airfoils the predicted target characteristics were met. Results for Riso-B1-18 showed a maximum lift coefficient of 1.64. A standa rd case of zigzag tape leading edge roughness caused a drop in maximum lift of only 3.7%. Cases of more severe roughness caused reductions in maximum lift between 12% and 27%. Results for the Riso-B1-24 airfoil showed a maximum lift coefficient of 1.62. The standard case leading edge roughness caused a drop in maximum lift of 7.4%. Vortex generators and Gurney flaps in combination could increase maximum lift up to 2.2 (32%). NOMENCLATURE

Modern wind energy industry is highly under the influence of the advances in aerospace and has borrowed and adopted many new and innovative concepts from that industry. With the aim of exploring potential benefits of flapping blades in wind turbines, this paper presents the preliminary results of the numerical simulation of a flapping-pitching rectangular flat plate in a uniform air flow. Various combinations of flapping amplitude, flapping frequency and pitching amplitude are analysed and their effect on the instantaneous and maximum lift coefficient is presented. The change in the flapping frequency and amplitude were shown to have considerable effect on the lift coefficient. It can be deduced from the results that the lift coefficient is influenced by the flapping frequency and flapping amplitude combination. For the size and geometry presented in this study, the lift coefficient is most affected by the flapping amplitude when compared to the flapping frequency. The results indicate that the pitching amplitude initially enhances the lift coefficient. However, excessive pitching amplitude results in low lift coefficient.

The roughness increase on horizontal axis wind turbine (HAWT) blade surface, especially on the leading edge, can lead to an aerodynamic performance degradation of blade and power output loss of HAWT, so roughness sensitivity is an important factor for the HAWT blade design. However, there is no criterion for evaluating roughness sensitivity of blade currently. In this paper, the performance influences of airfoil aerodynamic parameters were analyzed by the blade element momentum (BEM) method and 1.5 MW wind turbine blade. It showed that airfoil lift coefficient was the key parameter to the power output and axial thrust of HAWT. Moreover, the evaluation indicators of roughness sensitivity for the different spanwise airfoils of the pitch-regulated HAWT blade were proposed. Those respectively were the lift-to-drag ratio and lift coefficient without feedback system, the maximum lift-to-drag ratio and design lift coefficient with feedback system for the airfoils at outboard section of blade, and lift coefficient without feedback, maximum lift coefficient with feedback for the airfoils at other sections under the pitch-fixed and variable-speed operation. It is not necessary to consider the roughness when HWAT can be regulated to the rated power output by the pitch-regulated and invariable-speed operation.

While aerodynamic optimization of wing geometry in cruise flight has been widely investigated, such modifications inevitably alter the aerodynamic characteristics of high-lift devices, which remain essential during take-off and landing regimes. In this context, the present study addresses the influence of flap slot geometry on the aerodynamic characteristics of single-slotted flap configurations. For this reason, a parametrization method is introduced, combining cubic Bézier curves for flap definition with local curvature parameters for the cove region of the airfoil. This approach ensures geometric continuity in the retracted configuration while enabling rigorous control of the deployed flap geometry. The parametrization was integrated into an automated CFD and gradient-based optimization framework, enabling the efficient exploration of six geometric parameters across multiple configurations. The analysis revealed that lip length and the curvature of the flap’s upper surface have the most significant impact on aerodynamic performance, influencing lift generation, flow attachment, and drag reduction. Optimized configurations achieved up to a 7% increase in maximum lift coefficient and a 5% reduction in drag relative to the baseline geometry. These results highlight the potential of precise geometric control of flap slots to enhance aerodynamic efficiency, particularly in low-speed regimes relevant to take-off and landing. The proposed methodology establishes a foundation for advanced parametrization and optimization strategies for multi-element airfoil configurations and next-generation high-lift systems.

The aerodynamic performance of a three-wheeled vehicle was analysed using a Computational Fluid Dynamics (CFD) approach to evaluate its drag coefficient (Cd) and lift coefficient (Cl) under varying flow and environmental conditions. The study investigates the effects of velocity, air temperature, and density on aerodynamic forces, providing insights into their influence on vehicle stability and efficiency. A parametric analysis was conducted over a range of velocities (1–100 m/s), temperatures (223–328 K), and air densities corresponding to different altitudes (-1000 m to 80,000 m). The results indicate a direct correlation between velocity and aerodynamic forces, with Cd and Cl increasing non-linearly at higher speeds. Conversely, temperature variations showed negligible impact on aerodynamic performance, while decreasing air density at higher altitudes significantly reduced drag and lift forces. Additionally, a grid independence test was performed to ensure numerical accuracy, confirming that mesh refinement beyond a critical threshold yielded minimal variation in results. The three-dimensional model of the vehicle was developed in Fusion 360, and CFD simulations were conducted using ANSYS to assess aerodynamic behavior. The findings highlight the potential for aerodynamic optimization in three-wheeled vehicle design, emphasizing the need for refined body shaping and flow management strategies to enhance efficiency and stability. The motivation behind selecting a three-wheeled vehicle lies in its growing adoption for urban mobility, given its lightweight structure, lower manufacturing costs, and energy efficiency. However, its aerodynamic characteristics remain a critical factor in ensuring stability, particularly at higher speeds. This study highlights the potential for aerodynamic optimization in three-wheeled vehicle design, emphasizing the need for refined body shaping and flow management strategies to enhance efficiency and stability. This study serves as a foundation for future aerodynamic improvements in alternative vehicle designs, contributing to advancements in energy-efficient transportation.

Pulsed jet actuators (PJAs) are one of the candidate technologies to be integrated in Fowler flaps to increase the maximum lift coefficient of transport aircraft in the landing configuration. The total system consists of the actuators plus sensors, a piping system to supply pressurized air and a (redundant) power and communication system to provide actuator control. In this paper, it is investigated what increase in the maximum lift coefficient is required to justify the added weight and power off-takes that accompany the integration of pulsed jet actuators. This is done by making an automated design process for the overall aircraft, the piping assembly system, and the electrical wiring interconnection system. These last two sub-systems rely on KBE techniques that automate dimensioning and performance evaluation. A test case is specified that encompasses the design of a typical single-aisle mid-range aircraft with and without the PJA system installed. It is concluded that the introduction of the PJA system requires at least an increase in maximum lift coefficient of 0.2 to justify the increase in system mass and power off-takes. Furthermore, it is shown that if the maximum lift coefficient increases with 0.4, only small reductions in maximum take-off weight (−0.3 %) and operating empty weight (−0.6 %) can be expected, while the total fuel burn remains virtually constant.

CFD approach to modelling, hydrodynamic analysis and motion characteristics of a laboratory underwater glider with experimental results

Force measurement and surface oil flow visualization were carried out in wind tunnel to investigate the effects of wing’s planform and flexibility on aerodynamics for flow over low-aspect ratio wings. It is found that the empirical formulas proposed by Polhamus can present remarkably accurate predictions before stall for both rigid and flexible wings. The spiral vortex is responsible for rigid and flexible wing abrupt drop of lift coefficient after stall when the sweep angle is below 35°. The maximum lift coefficient and the stall angle for flexible wings are smaller than those for rigid ones when the sweep angle is below 40°, and the wing flexibility nearly does not influence the aerodynamics as the sweep angle is beyond 40°. Moreover, lift coefficient curves exist hysteresis loop in the following range of sweep angle: 10°–35° for the flexible wings and 5°–50° for the rigid wings. It can be deduced from oil visualization experiment that the sudden disappearance of spiral vortex flow structure may be resp...

The flow field of a NACA 0024 airfoil is investigated using ANSYS CFX 14. The airfoil used is equipped with a rotating cylinder mounted at its leading edge. Shear Stress Transport (SST) k-ω turbulence model is used in accordance with time dependent simulations. The numerical results are compared with published experimental data. The mean lift coefficient, the pressure coefficient and the velocity profiles at different sections of the airfoil surface are obtained and compared with Experimental results. Both cases of rotation and no rotation of the cylinder are presented. The maximum lift coefficient as well as the stall angle is increased when the cylinder rotation is present. The numerical data of lift coefficient values compared well with the experimental data.

Introduction I wing design it is very important to determine the maximum lift coefficient as accurately as possible, since this lift coefficient corresponds to the stall speed that is the minimum speed at which controllable flight can be maintained. Any further increase in angle of incidence will increase flow separation on the wing upper surface, resulting in a loss of lift and a large increase in drag. Despite the significant advances in computational fluid dynamics, our ability to predict the maximum lift coefficient and poststall behavior of single airfoils, wings, multielement airfoils, or wings is still not satisfactory. The chief culprit is the lack of an accurate turbulence model to represent flows with extensive separation. Inaccuracies of numerical solutions of the conservation equations at these flow conditions, as well as difficulties in modeling of flow near the trailing edge of an airfoil or wing, add to this dilemma. Recently Cebeci et al. described an interactive boundary-layer method, together with the e approach to the calculation of transition, for predicting stall and poststall behavior of airfoils at low and high Reynolds numbers. The turbulence model was based on the Cebeci-Smith algebraic eddy-viscosity formulation with improvements for strong pressure gradient effects and transitional flows at low Reynolds numbers. Comparison of calculated results for incompressible flows indicated good agreement with experiment for a wide range of Reynolds numbers. Preliminary calculations for low-Mach-number flows with this interactive method that applies compressibility corrections to the panel method also indicated good agreement with data and showed that at a Mach number of 0.30, the compressibility effect on (Q)max was not negligible. The ability of the calculation method of Ref. 1 to predict the effect of ice on airfoil stall at high Reynolds numbers is examined here, and the results are presented in the following section. The interactive boundary-layer method previously developed for iced airfoils and coupled to the LEWICE code for predicting ice shapes' is replaced with the calculation method of Ref. 1. Appropriate changes are made to this method, including roughness effects due to ice, and the results are presented in the following section.