Leading-edge Slat Research Articles

Numerical Study on Single and Multi-Element NACA 43018 Wing Airfoil with Leading-Edge Slat and Slotted Flap

An optimal wing configuration is crucial for achieving the best performance during various flight phases, including take-off, cruising, and landing. Such configurations also contribute to maximizing the aircraft's cruising range. This study compares the aerodynamic performance of NACA 43018 wings under different conditions: without high-lift devices, with a slotted flap, and with a combination of a leading-edge slat and slotted flap. Numerical simulations were conducted using the k-ɛ Realizable turbulence model at twelve different angles of attack, with a flow speed of 120 m/s. The results demonstrate that multiple-element wings significantly improve aerodynamic performance, particularly at low angles of attack, by reducing the induced drag coefficient and delaying flow separation.

Impact of leading edge slat on stall performance of multi-element airfoils

Impact of leading edge slat on stall performance of multi-element airfoils

Experimental investigation of suction and oscillatory blowing as an alternative for a leading edge slat

High lift devices such as flaps and slats have conventionally served in multi-element wings to enhance aerodynamic performances in general, and lift in particular. However, more innovative solutions are required to address the evolving challenges the aviation industry faces to reach higher lift capabilities. Active flow control (AFC) has been studied for many years but applications have remained strictly experimental and did not reach any widespread industrial use to enhance aerodynamic performance. This experimental study investigates applying a novel AFC concept to a high-lift wing configuration to replace the slat with AFC actuators that incorporate suction and oscillatory blowing (SaOB) into the main element of a thick airfoil. The multi-element wing is equipped with an AFC system to investigate the possibility of replacing the leading edge slat with an AFC actuator. The benefits of having an AFC system acting as a virtual slat are discussed quantitatively, with experimental results, and practically as an efficient solution for designing a lighter and more capable high-lift device. This research aims to characterize the application of AFC to a wing in terms of aerodynamics and energy efficiency.

Investigation on Solid Particle Erosion Performance of Aluminum Alloy Materials for Leading-Edge Slat

This study aims to characterize solid particle erosion behaviors of three different aluminum alloys (AA2024-T351, AA6061-T651, and AA7075-T651) and reveal their erosion performances on leading-edge slat of airplane wings. Solid particle erosion tests were conducted using silicon carbide erodent particles under the conditions of six different impingement angles (20°-90°) and four different impact velocities (70-192 m/s). The erosion simulations of a leading-edge slat of the aforementioned aluminum alloys were numerically simulated at four different rotation angles (0°-15°) for three different impact velocities (130-250 m/s). A commercial ANSYS Fluent software using the Euler-Lagrange equation and an experimental data-based erosion model was used for the erosion simulations. The experimental results showed that the erosion rate increases with increasing impact velocity and the maximum erosion rate is obtained at the impingement angle of 30° which reflects the ductile manner. Of the three aluminum alloys, the AA6061-T651 exhibited the worst erosion behavior followed by the AA2024-T351 sample, whereas the AA7075-T651 had the best erosion resistance. The numerical results indicated that the erosion rate values of slat surfaces made up of three different aluminum alloys showed a slight increase after a slat rotation angle of 5°.

Morphing concepts in the field of rotorcraft

Drawing inspiration from avian creatures, aeronautical engineers strive to create an ideal wing design that can seamlessly perform in various flight conditions. Avian creatures, as well as bats and other flying organisms, exhibit a striking aptitude to adjust the lift produced by their wings, displaying the capacity to repeatedly tailor their wing configurations to match specific environmental conditions. An example of this is when their wings are tightly tucked during dives for hunting, or fully extended during gliding to conserve energy. Moreover, these organisms can manipulate the curvature and twist of their wings to maintain precise control over their aerial maneuvers. In contrast, engineers in the aircraft industry continue to rely on the standard, robust and structured “one-point design” approach, which remains the most practical and feasible method to apply. Nonetheless, advancements in technology have emerged to address long-standing challenges in wing manufacturing that were previously deemed insurmountable. This convergence of different technologies has given significant momentum and recognition to the field of “morphing discipline”. When considering an aircraft, shape changes primarily relate to the wing of a fixed-wing aircraft or the blade of a rotorcraft. The concept of achieving a "smooth" shape change stems from the crucial need for drag reduction and improved flow quality, resulting in improved overall performance. The state-of-the-art morphing concepts applied to rotorcrafts comprise a wide range of investigations aimed at improving performance. Looking ahead, the primary challenge for morphing technology will be to persuade the industry of its tangible benefits. This encompasses enhanced aerodynamic efficiency, minimized installation footprint when contrasted with conventional control surface mechanisms, reduced overall weight, and an equivalent standard of safety. This research provides an overview of the current development of different control devices and explores the impact of previous and continuous research endeavors in this field. Numerous ideas for managing airflow have been explored with the aim of enhancing the performance abilities of rotary-wing aircraft. These include active morphing in rotorcraft such as leading edge slats, trailing edge flaps, and passive morphing in rotorcraft such as variable rpm rotorcraft. The aim of these blade modifications is to achieve various desired effects such as increasing the maximum lift coefficient, reducing drag, and minimizing vibratory loads. Convincing the industry of these advantages will play a crucial role in shaping the future of morphing technology.

Noise reduction for high-lift devices on a swept wing model by droop nose

The noise of slat in the high-lift device of large aircraft is one of the main noise sources of large aircraft during take-off and landing. This paper investigates a noise reduction measure which consists in droop nose for high-lift device of three-dimensional sweep wing instead of leading-edge slat. The unsteady flow field and the far-field noise are predicted by a hybrid SA-IDDES method coupled with the integration of the Ffowcs Williams–Hawkings equation (FW-H). Based on the optimal droop nose configuration from 2D airfoil, the results of the 3D swept wing model with the droop nose on the leading edge of the inner wing are given. Numerical results show that the droop nose configuration can reduce the overall sound pressure level (OASPL) of the intermediate frequency noise, and the peak noise can be reduced by 15 dB, and the sound pressure level of the broadband noise at low and high frequencies is also partially suppressed. The maximum lift-to-drag ratio of the droop nose configuration increased by 2.7 %. In addition, the noise reduction mechanism of leading-edge droop is analyzed by combining the vortex sound theory and the generation and evolution process of the vortex structure. The leading-edge droop nose configuration fundamentally not only eliminates the vortex structure of the slat cove of the leading-edge slat configuration, but also avoids the interaction between the slat slot flow and the upper surface of the main wing. These directly lead to a thinner boundary layer vortex near the top of the main wing, and the wall pressure pulsation is smaller than that of the leading-edge slat high-lift device, thereby reducing noise. The reduction of the BEF distribution on the upper surface boundary shows again that the droop nose inhibits the vortex on the upper surface of the main wing. The change of the Lamb vector divergence reveals the noise reduction mechanism of the drooping leading edge configuration directly from the sound source.

Experimental Study of Far-Field Radiated Noise Characteristics for the Leading-Edge Slat of a 30P30N Airfoil

Experimental Study of Far-Field Radiated Noise Characteristics for the Leading-Edge Slat of a 30P30N Airfoil

Parametric Investigation of Leading-Edge Slats on a Blended-Wing-Body UAV Using the Taguchi Method

The current study investigated the effect of leading-edge slats on the longitudinal stability at high angles of attack of a Blended-Wing-Body (BWB) Unmanned Air Vehicle (UAV). Using a Design of Experiments (DOE) approach and, specifically, the Taguchi method, four leading-edge slat design parameters were investigated on three different levels. These parameters were the slat semi-span, the rotation of the slat element, the extension forward of the leading edge and the downward drop below the leading edge. An L9 orthogonal array (OA) was used to investigate the influence of these key design parameters using three performance criteria, namely the angle at which pitch break occurs, the corresponding speed and the distance between the Neutral point of each configuration and the Neutral point of the reference platform. The investigation was conducted by using high-fidelity Computational Fluid Dynamics (CFD) methods for each of the nine configurations defined by the L9 OA, over a range of angles of attack between −4 and 16 degrees. Based on these results, and using a Signal-to-Noise ratio (SNR) analysis, two combinations were eventually derived, one that optimized pitch break angle and speed and one that optimized longitudinal stability. Finally, the Pareto Analysis of Variance (ANOVA) technique was conducted to define the contribution of each of the six design parameters on the selected performance criteria. More specifically, the semi-span seemed to have the most significant effect on pitch break angle and speed, whereas the rotation of the slat element was the most important parameter with regard to static stability.

Characterization of Slat Noise Radiation from a High-Lift Common Research Model

A test campaign was conducted in the open-jet test section of the NASA Langley Research Center’s 14- by 22-Foot Subsonic Tunnel on a 10%-scale semispan version of the High-Lift Common Research Model incorporating a leading-edge slat, a trailing-edge flap, and removable high-fidelity main landing gear. Computational simulations were used to support the model development, provide important information such as load distributions along the high-lift elements, and aid in the design of noise reduction devices. Aerodynamic and far-field acoustic measurements were obtained on baseline and acoustically treated model configurations using an ensemble of time-averaged and unsteady surface pressure sensors paired with a traversing 97-microphone phased array viewing the pressure side of the airframe. A deconvolution method incorporating corrections for shear layer acoustic wave decorrelation was employed to determine the locations and strengths of relevant noise sources along the span of the slat and in the vicinity of the flap edges and landing gear. The main goal of the test campaign was to use the surface pressure and array data to evaluate the effectiveness of various slat noise reduction concepts. It was found that slat-gap fillers produced significant broadband noise reduction with minimal aerodynamic performance penalties for all Mach numbers and airfoil angles of attack that were investigated. The test campaign clearly demonstrated the noise reduction benefits that can be obtained by applying appropriate treatments to leading-edge slats on commercial transport-class aircraft.

A New Flow Control Method of Slat-Grid Channel-Coupled Configuration on High-Lift Device

A slot formed between the slat and the main wing of the 2D high-lift device is used to accelerate the convergence of the flow to the upper surface of the main wing to improve the flow field quality. In order to further enhance flow characteristics, this paper proposes a design method for multi-channel leading-edge slats based on grid flow channels. On the one hand, a specific shape of a shrinking expansion tube is formed to improve the lift characteristics of the leading slat. On the other hand, the newly formed slot plays a similar role to that of the jet stream, delaying the separation on the upper surface of the main wing, making the separation point move back and helping to improve the lift characteristics of the main wing. The optimization of coupled slat-grid channel configuration is developed by using the DOE algorithm. The geometric parameters, such as coordinates and curve slope, are considered as design variables, and the maximum lift–drag ratio is taken as the optimization objective required to obtain the optimal configuration. The simulation and optimization results show that the lift coefficient increases by 3.3%, the drag coefficient decreases by 12.7%, and the lift–drag ratio increases by 18.4% of the optimal configuration compared with the original airfoil at an angle of attack of 16.3°.

Effect of Leading-Edge Gap Size on Multiple-Element Wing NACA 43018

Generally, large airplanes are added with more than one high lift device to improve aerodynamic performance. Each configuration of the addition of high lift devices will result in different aerodynamic performances. The added high-lift devices include flaps, leading-edge slats, vortex generators, fences, and others. This study shows the aerodynamic performances of the NACA 43018 wing airfoil with the addition of leading-edge slats and plain flaps in cruise conditions with variations of leading-edge gap size. The angles of attack used include 0°, 2°, 4°, 6°, 8°, 10°, 12°, 15°, 16°, 17°, 19°, and 20°. Numerical simulations have been performed using Ansys Fluent 19.1 with a k-ε Realizable turbulent model. The aerodynamic performance of the configuration that uses the leading-edge slat shows a higher value than the rectangular wing even though the lift and drag coefficients produce lower values. The use of leading-edge slats produces fluctuations, which are indicated by different values of the pressure coefficient, that show the interaction of adding momentum from the main flow from the direction of the leading edge. The increase in the angle of attack causes bubble separation in the leading-edge gap size, which is like a pseudo-body transforming from convergent to smaller.

Numerical Analysis of Slotted Wings Using Fluid-Structure Interaction

For shorter landing and take-off path in airports, the aircrafts should reduce their speed with keeping high lifting force. This paper is to identify solutions to increase the lift force of the wing significantly under several flight scenarios (such as takeoff and landing) using leading-edge slats and their relationship with the dynamic parameters of the aerodynamic wing. The study is performed by the use of ABAQUS 2016 software. The problem is solved for turbulent flow and 2-dimensional composite wing at constant Reynolds’s number of (6.49 × 105) and constant boundary conditions. Various depths have been used for the auxiliary airfoil at constant width and gap. All stresses at the wing base were obtained. The pressure distribution on the airfoil surface was determined, air velocity distribution was tracked over the surface, lift and drag forces and their coefficients were computed. The results show that the highest value of the lift coefficient is 0.489 at the depth (-3 %) of the wing chord, it decreases when the depth of the slat becomes zero %, and the rise returns with increasing depth to (4 %), but it does not reach the maximum value, while the highest drag coefficient was (1.89) at depth (4 %) of the wing chord. The maximum value of Von Mises stress was found at depth of 4 % with value of 1.605 × 105 Pa.

Bio-inspired optimization of leading edge slat

PurposeThe high lift devices are effective at high angle of attack to increase the coefficient of lift by increasing the camber. But it affects the low angle of attack aerodynamic performance by increasing the drag. Hence, they have made as a movable device to deploy only at high angles of attack, which increases the design and installation complexities. This study aims to focus on the comparison of aerodynamic efficiency of different conventional leading edge (LE) slat configurations with simple fixed bioinspired slat design.Design/methodology/approachThis research analyzes the effect of LE slat on aerodynamic performance of CLARK Y airfoil at low and high angles of attack. Different geometrical parameters such as slat chord, cutoff, gap, width and depth of LE slat have been considered for the analysis.FindingsIt has been found that the LE slat configuration with slat chord 30% of airfoil chord, forward extension 8% of chord, dip 3% of chord and gap 0.75% of chord gives higher aerodynamic efficiency (Cl/Cd) than other LE slat configurations, but it affects the low angles of attack aerodynamic performance with the deployed condition. Hence, this optimum slat configuration is further modified by closing the gap between LE slat and the main airfoil, which is inspired by the marine mammal’s nose. Thus increases the coefficient of lift at high angles of attack due to better acceleration over the airfoil nose and as well enhances the aerodynamic efficiency at low angles of attack.Research limitations/implicationsThe two-dimensional computational analysis has been done for different LE slat’s geometrical parameters at low subsonic speed.Practical implicationsThis bio-inspired nose design improves aerodynamic performance and increases the structural strength of aircraft wing compared to the conventional LE slat. This fixed design avoids the complex design and installation difficulties of conventional movable slats.Social implicationsThe findings will have significant impact on the fields of aircraft wing and wind turbine designs, which reduces the design and manufacturing complexities.Originality/valueDifferent conventional slat configurations have been analyzed and compared with a simple fixed bioinspired slat nose design at low subsonic speed.

Slat noise measurements in open-jet, hard-wall and hybrid wind tunnel test sections

This paper assesses the comparability of aeroacoustic slat noise measurement in an open-jet, a hard-wall, and a hybrid test section configuration. The 30P30N common research model was tested in the Aeroacoustic Wind Tunnel of the University of Twente at a chord-based Reynolds number up to 1⋅106 and free-stream Mach number 0.15. Pressure distribution measurements were used to determine the aerodynamic similar condition of the flow in the slat cove. These conditions are subsequently validated by time-averaged 2D PIV measurements. The experiments show that general aspects of the flow in the slat cove are similar when the Cp distribution around the leading-edge slat element in each test section configuration is comparable. Overall, the shear layer path and reattachment point are in close agreement, showing that the aeroacoustic noise production mechanisms are comparable. Microphone phased array measurements were subsequently conducted to compare the far-field noise characteristics originating from the slat in each test section configuration. A framework is proposed to correct the acoustic measurements of each test section configuration to a standard free-field condition. Transmission loss corrections are required for the microphone measurements in the hard-wall and hybrid test section configurations, whereas a coherence loss model was validated and applied to correct the microphone measurements in the open-jet. The measured noise spectra in each test section configuration show differences up to approximately 5 dB. In general the differences were found to be smaller. The largest differences are seen for small angles of attack with noise levels in the open-jet test section showing considerably higher values than the hard-wall test section and hybrid test section results. In addition, the behavior of the vortex shedding hump in the high frequency range is different in the open-jet measurements, which could be related to the larger noise level differences at small angles of attack.

Airfoil Morphed Leading-Edge Design for High-Lift Applications Using Genetic Algorithm

A genetic algorithm was developed for the optimal design of morphing airfoil leading-edge shapes to maximize airfoil , while retaining a constant leading-edge arc length. Using a candidate slotted, natural laminar flow airfoil, an inviscid–viscous flow simulation program suite was used to evaluate a randomized population of airfoils during the design process at low-speed, high-lift flight conditions of a representative single-aisle commercial transport aircraft. The use of a continuous morphing strategy to perform treatment of the leading-edge geometry in high-lift scenarios, as opposed to leading-edge flap or slat elements, is envisaged as a means to mitigate surface discontinuities and retain laminar flow qualities of future airfoil sections in cruise conditions. By using different definitions of the objective cost function, multiple morphed airfoil designs were found through the genetic algorithm to create a library of possible designs aimed to purposefully alleviate the leading-edge stall tendencies of the baseline airfoil. Three morphed designs were validated through experimentation in a low-speed, open-return wind tunnel. Resulting data were used to verify that the leading-edge separation property of the baseline airfoil during stall was effectively alleviated with use of the morphed leading-edge geometries.

Effect of the trailing-edge flap on tones due to self-excited oscillation within the leading-edge slat cove

To study the tonal noise characteristics of the leading-edge slat of high-lift configurations with and without a deployed trailing-edge flap, experiments are conducted in the D5 aero-acoustic wind tunnel at Beihang University. The numerical simulation method is used to obtain the necessary flow information. The experimental results show that low to mid frequency tonal noise generated by self-excited oscillation within the slat cove is dominant and its corresponding frequencies are basically unchanged whether the flap is deployed or not. However, the primary mode of self-excited oscillation within the slat cove switches to a higher one when the flap is deployed. Further analysis results demonstrate that variation of the primary mode is found to be closely related to the flow characteristics in the self-excited oscillation feedback loop. The number of the primary mode is generally proportional to the ratio between the vortex shedding frequency and the self-excited oscillation frequency. The flap being deployed results in an increase in the effective angle of attack of both the main wing and slat, which leads to a thinner separating boundary layer, thus increasing further the vortex shedding frequency and this ratio.

Noise characteristic of the nacelle/pylon/slat juncture in a realistic high-lift aircraft configuration

The leading-edge slat is an important contributor to airframe noise. Recent studies revealed that the leading-edge slat in realistic high-lift aircraft configurations forms a more complex nacelle/pylon vortex system, resulting in an intense noise source. Thus, this study investigated the noise characteristics of leading-edge slats in realistic aircraft configurations based on phased microphone arrays. The non-negative L1/2 regularization method for the deconvolution approach for the mapping of acoustic sources (DAMAS) was applied to the airframe noise benchmark test DLR1. An uneven and concentrated source strength distribution around the nacelle/pylon/slat juncture region was obtained owing to the three-dimensional flow pattern. Thus, the leading-edge region was partitioned into small sub-regions to extract the nacelle/pylon/slat juncture noise source. The angle of attack (AoA) has different effects on the slat and nacelle/pylon/slat juncture noises. In detail, the minimum amplitude of the slat noise was obtained at AoA = 5°. In contrast, the amplitude of the nacelle/pylon/slat juncture noise increased as the AoA increased, especially at a high AoAs. The significant increase in the amplitude of the nacelle/pylon/slat juncture noise was mainly attributed to the large increase in the spectrum at the low-frequency range with the increase in the AoA. In addition, the amplitude levels of the slat and nacelle/pylon/slat juncture noises approximately follow the fifth power law of the Mach number with their spectra following a similar scaling law. Both noise spectra followed the Strouhal number scaling laws in the low-frequency range and Helmholtz number scaling law in the high-frequency range. However, both noise sources have different noise generation mechanisms. The revealed noise characteristics are helpful for establishing a physics-based nacelle/pylon/slat juncture noise prediction model and subsequently, the development of noise reduction technologies.

Aerodynamic characteristic of wind turbine with the leading edge slat and Microtab

The leading edge slat can increase the flow energy of airfoil suction surface, and Microtab can hinder the flow of pressure surface. In order to combine the advantages of two passive flow control methods, this paper creatively proposed to combine the leading edge slat with Microtab to simulate the S809 airfoil, which slat had the contour characteristics of S809 airfoil. The transition SST turbulence model was applied to estimate the aerodynamic characteristics of the S809 airfoil. According to research, the flow control approach of the leading edge slat combined with Microtab could effectively restrain the flow separation and ameliorate the aerodynamic characteristics of the S809 airfoil. For Case-12, when the airfoil angle of attack was 16.22°, the boundary layer flow separation was restrained completely, Cl was increased to 3.06 from 1.79 and increased by 171% compared with clear airfoil. Research shown that the position of slat, the position and height of the Microtab were significant factors for the aerodynamic characteristics of the S809 airfoil. Microtab position at 0.95c and height at 0.03c performed the best effect to inhibit flow separation and improve lift coefficients.

Effects of leading edge slat on the aerodynamic performance of low Reynolds number horizontal axis wind turbine

The present study considers the use of a slat near the leading edge of an airfoil to delay or completely cancel separation and thus improves the performance of small wind turbines, in addition to lowering the cost of the blade by reducing the used material. Flow around the S809 airfoil with a slat is numerically simulated using ANSYS 2019 R1 (CFD) program. The present work concentrates on two geometrical parameters namely the location of the slat relative to the base airfoil and the chord size of the slat. Relative to the total chord length of the main airfoil, a slat with a chord length 10% is located at range of distances +3.6% to +10.8% in y direction, slat located at 9% shows the best lift coefficient. To study the effect of slat size, slat sizes range from 5% to 12.5% are numerically investigated, slat size 7.5% shows the best lift coefficient and attached flow over the airfoil for a wide range of angles of attack. The improvement is achieved with the same total chord length with 12% saving of material which corresponds to 6% of the total cost of the turbine.

Flow Parameters in Dominant Mode Selection of Self-Excited Oscillation in Slat Coves

This paper investigates self-excited oscillation within a slat cove and the flowfield parameters involved in selection of the main oscillation mode. Aeroacoustic experiments are described for aerodynamic noise characteristics on the leading-edge slat of a two-dimensional three-element high-lift configuration. Different geometrical settings are tested at varying angles of attack. A numerical simulation method is used as an auxiliary tool to analyze the flowfield. Based on previous research results and the experimental results obtained in this paper, a ratio of is proposed to evaluate the main mode of self-excited oscillation, where is the shear-layer length within the slat cove, is the boundary-layer thickness near the point of separation, and is the ratio of average vortex convection velocity along the shear layer to the maximum velocity in the shear layer near the cusp. The relationship between the aforementioned ratio and the main mode of oscillation is verified by far-field noise data for a high-lift configuration with different geometrical settings. As the slat translation along the chordwise direction of the main wing increases, the ratio also increases; and the main mode of self-excited oscillation switches from mode II to a transition phase, and then to mode III.