Low Angles Of Attack Research Articles

Computational analysis of SD7037 airfoil with plain flap

The impact of a simple trailing-edge plain flap on the aerodynamics of the SD7037 airfoil have been studied in this paper using computational fluid dynamics at Reynolds number of 3×105 across various low angles of attack and flap deflection angles. The computational model was evaluated by using Star CCM+ software with κ--ω SST turbulence and gamma transition model to solve Navier-Stokes equations. The accuracy of the computational model has been confirmed through comparison with experimental data, showing a high level of agreement at low angles of attack. The findings revealed that specific combinations of angles of attack and flap deflection angles could increase the lift-to-drag ratio by over 70% compared to baseline conditions, benefiting airfoil performance, particularly during takeoff. Some combinations, however, resulted in decreased performance and should be avoided. The results also showed that with the increase of either the angle of attack or the flap deflection angle, the pitching moment increased.

Proper orthogonal decomposition analysis for two-dimensional wake transition behind a NACA 0012 airfoil in a soap film

Abstract An experimental study has been conducted to analyze the flow characteristics around a NACA 0012 airfoil in a horizontal soap film tunnel. The investigation involves both the qualitative observations and the quantitative measurements at various angles of attack (0° to 20°) for Reynolds numbers ∼ 3000 and 5000. Flow visualizations were carried out using the interference technique, while particle image velocimetry (PIV) was used for velocity field measurements. The study examines the two-dimensional wake transition of the NACA 0012 airfoil under different Reynolds numbers and angles of attack in a soap film tunnel. Results indicate that changes in the trailing flow within the soap film are primarily influenced by the Reynolds number and angle of attack. The visualization of the flow exposes different states of wakes and their corresponding flow structures. At lower angles of attack α ≤ 6°, vortex shedding occurs in an alternating mode, while at intermediate angles (8° to 16°), the wake widens. The vortices exhibit more chaotic and turbulent behavior at higher angles of attack, i.e., α ≥ 16°. Furthermore, the study indicates that the same wake transition takes place at lower angles of attack as Reynolds numbers rise.

Accuracy and Computational Cost Comparison of Numerical Flight Dynamics Reduced-Order Models

This paper compares several approaches to model the unsteady and nonlinear longitudinal aerodynamics of a generic unmanned combat air vehicle configuration based on computational fluid dynamics (CFD). Three different reduced-order models (ROMs) are implemented through unsteady simulations to predict the evolution of forces and moments during prescribed trajectories carried out at M=0.15. The first model investigates a linear quasi-steady model based on a first-order Taylor series expansion of the aerodynamic coefficients. The second one picks up the mathematical formulation of the first model but adds a dynamic coefficient in order to take into account the unsteady effect of rotation rate variations. Finally, the third model is built on indicial functions computed from positive step changes in the angle of attack or pitch rate. The unknown parameters of the first two models and the unknown indicial functions of the third one are obtained, respectively, from the study of forced oscillations and step motions. These prescribed trajectories are performed using unsteady Reynolds-averaged Navier–Stokes simulations coupled with a grid displacement tool. Two different methods are tested and compared to carry out the CFD mesh motions: an overset method and a full moving grid (FMG) method. The indicial functions obtained using the FMG method give results equivalent to those found in the literature in half the time of the overset method. To reduce the computational costs to set up the ROMs, the FMG method is preferred to the overset method. ROMs are then applied, and their predictions are compared with CFD simulations for prescribed trajectories with different angular rates and angles of attack. A study of instantaneous and overall accuracy associated with the implementation cost of the ROMs has allowed to show their strengths and their weaknesses. Similar results are observed at low angles of attack for the ROMs studied, while a major modeling advantage of the indicial method has been identified for higher angles of attack in the nonlinear domain.

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.

Parametric Optimization Study of Novel Winglets for Transonic Aircraft Wings

This paper deals with the topic of reducing drag force acting on aircraft wings by incorporating novel winglet designs, such as multi-tip, bird-type, and twisted. The high-speed NASA common research model (CRM) was selected as the baseline model, and winglet designs were retrofitted while keeping the projected wingspan constant. Computational analysis was performed using RANS coupled with the Spalart–Allmaras turbulence model to determine aerodynamic coefficients, such as CL and CD. It was observed that the multi-tip and bird-type designs performed exceptionally well at a low angle of attack (0°). A parametric study was conducted on multi-tip winglets by tweaking the parameters such as sweep angle (Λ), tip twist (Є), taper ratio (λ), and cant angle (Φ). The best combination of parameters for optimal aerodynamic performance while maintaining the wing root bending moment was determined using both the Taguchi method and Taguchi-based grey relational analysis (T-GRA) coupled with principal component analysis (PCA). Also, the percentage contribution of each parameter was determined by using the analysis of variance (ANOVA) method. At the design point, the optimized winglet design outperformed the baseline design by 18.29% in the Taguchi method and by 20.77% in the T-GRA coupled with the PCA method based on aerodynamic efficiency and wing root bending moment.

Aerodynamic flow field analysis of vortex interactions on multi-swept wing configurations

The underlying flow physics of vortex breakdown and vortex-vortex interactions at moderate to high angles of attack is not well understood. This lack of understanding is a challenge to the development of optimized wing designs for fighter aircraft. The goal of this study is to gain insight into the breakdown of vortices on multi-swept wing configurations. Two generic models were considered, one with a gradual vortex breakdown and one with a rapid (abrupt) breakdown. Aerodynamic forces were calculated from the US Air Force Academy (USAFA) subsonic wind tunnel to validate the experimental setup using computational fluid dynamics (CFD) and previous work. Stereoscopic particle imaging velocimetry (SPIV) was used to visualize vortex-vortex interactions, measure vortex strengths, and identify vortex breakdown locations. The measured lift and drag forces showed flow physics that agreed with previous experiments and CFD, supporting the validity of the new experimental setup. Analysis of SPIV and CFD data improved our understanding of the different types of vortex interactions observed for both models. The only difference between the models was the location of the secondary sweep angle on the chord. The C4L80 model, with a secondary sweep angle at 40% of the root chord, had a large shear layer along the leading edge, which led to vortex merging and gradual vortex breakdown at lower angles of attack than its counterpart. The C6L80 model, with a secondary sweep angle at 60% of the root chord, had no such shear layer and instead developed strong, isolated vortices that were characterized by vortex braiding and abrupt breakdown at high angles of attack. These vortices were entrained into the main vortex at the leading edge of the strake (inboard vortex), creating a balanced vortex system that enhanced lift and delayed the onset of gradual vortex breakdown seen in the C4L80 model.

Experimental Investigation on the Aerodynamic Parameters of Trailing Edge Wake Generators

In the realm of aerodynamics, the investigation of airfoil performance stands as a critical domain, with an ever-growing emphasis on optimizing designs for enhanced efficiency. This study investigates the aerodynamic performance of a NACA 0015 airfoil featuring various trailing edges, including serration, comb, poro-serration, and comb-serration. The experiments were conducted in a wind tunnel at angles of attack ranging from -15° to 15° and Reynolds numbers of 1.5 x 105 and 2.0 x 105. To accurately quantify the forces and moments acting on the airfoil models, a calibrated six-component balance was utilized to measure the aerodynamic coefficients of each airfoil configuration. The lift coefficient (cl), drag coefficient (cd), and pitching moment coefficient (cm) are analyzed for the various trailing edge types and Reynolds numbers. Results indicate that the baseline model demonstrated better aerodynamic performance compared to other types of trailing edge. Most trailing edges, except the baseline, resulted in a decrease in the lift coefficient. However, at very low angles of attack, the airfoil showed an improvement in the maximum lift coefficient. Most trailing edges exhibited an increase in the drag coefficient at a Reynolds number of 1.5 x 105. However, at a Reynolds number of 2.0 x 105, the drag coefficient showed a similar trend as the baseline. All types of trailing edges, including the baseline, displayed a similar trend in the pitching moment coefficient. When evaluating the lift coefficient to drag coefficient ratio, all trailing edges generally performed similarly at all Reynolds numbers. In general, the baseline model emerges as the optimal choice, showcasing superior aerodynamic characteristics across the evaluated parameters.

Effects of structural geometric nonlinearities on the transonic aeroelastic characteristics of wing

The structural geometric nonlinearity has a great influence on the transonic aeroelastic characteristics of aircraft wings, which is an important problem for the aerodynamic and structural design of wing. In this paper, a CFD/CSD coupling method based on the high precision in-house solver is first established, and the reliability of the CFD and the CFD/CSD coupling method is then verified. Subsequently, the effects of geometric nonlinearity on the transonic aeroelastic characteristics of wing under different dynamic pressures are studied. The results show that: (1) At low dynamic pressure conditions, for geometric linear wing structure, limit-cycle oscillation occurs at various angles of attack, and the limit-cycle oscillation amplitude decreases with the increases of angle of attack. At low angle of attack, the wing deformation characteristics are dominated by first bending and first torsion, and the first torsion weaken as the angle of attack increases. For the geometric nonlinear wing structure, the limit-cycle oscillation occurs only at an attack angle of 0° As the angle of attack increases, the structural displacement eventually exhibits static deformation characteristics. (2) Under high dynamic pressure conditions: the displacement of geometric linear wing structure presents a divergent trend. For geometric nonlinear wing structure, the limit-cycle oscillation amplitude decreases first and then increases with the increase of the angle of attack. At low angle of attack, the deformation characteristics of the wing are dominated by first bending and first torsion. With the increase of the angle of attack, the first torsion weakens, while the second bending strengthen. With the further increase of the angle of attack, the second bending weaken and become co-dominated by the first bending and first torsion. It can be seen from the above results that the aeroelastic characteristics of both the geometric linear structure and the geometric nonlinear structure wings are significantly different. These findings provide valuable insights for the meticulous design of aircraft wings.

Resolvent model for aeroacoustics of trailing edge noise

This study presents a physics-based, low-order model for the trailing edge (TE) noise generated by an airfoil at low angle of attack. The approach employs incompressible resolvent analysis of the mean flow to extract relevant spanwise-coherent structures in the transitional boundary layer and near wake. These structures are integrated into Curle’s solution to Lighthill’s acoustic analogy to obtain the scattered acoustic field. The model has the advantage of predicting surface pressure fluctuations from first principles, avoiding reliance on empirical models, but with a free amplitude set by simulation data. The model is evaluated for the transitional flow (Re=5e4\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\ ext {Re} = 5e4$$\\end{document}) around a NACA0012 airfoil at 3 deg angle of attack, which features TE noise with multiple tones. The mean flow is obtained from a compressible large eddy simulation, and spectral proper orthogonal decomposition (SPOD) is employed to extract the main hydrodynamic and acoustic features of the flow. Comparisons between resolvent and SPOD demonstrate that the physics-based model accurately captures the leading coherent structures at the main tones’ frequencies, resulting in a good agreement of the reconstructed acoustic power with that of the SPOD (within 4 dB). Discrepancies are observed at high frequencies, likely linked to nonlinearities that are not considered in the resolvent analysis. The model’s directivity aligns well with the data at low Helmholtz numbers, but it fails at high frequencies where the back-scattered pressure plays a significant role in directivity. This modeling approach opens the way for efficient optimization of airfoil shapes in combination with low-fidelity mean flow solvers to reduce TE noise.Graphical abstract

Wing design optimization and stall analysis with Co-flow Jet active control

Coupled with Co-flow Jet (CFJ) technology, the Non-dominated Sorting Genetic Algorithm II was utilized for the multi-objective combination optimization of an optimized Co-flow Jet wing, based on National Advisory Committee for Aeronautics (NACA) 6421. A high-precision numerical simulation using the delayed detached eddy simulation model was performed on the optimized wing to investigate the three-dimensional flow separation characteristics after static stall. The stall improvement was investigated by adjusting the momentum coefficient of the injection. The results show that the optimized wing exhibits significant improvements in aerodynamic performance and corrected aerodynamic efficiency. At an angle of attack of 10°, the average lift increased by 16.25% and the drag decreased by 27.23% compared to the CFJ6421 wing, while effectively addressing the problem of low modified aerodynamic efficiency of the CFJ wing at lower angles of attack. By utilizing higher momentum and improving the boundary layer control capability, flow separation is effectively suppressed, thus achieving the goal of stall recovery of the CFJ wing.

Effects on Aircraft Performance Due to Geometrical Twist of Wing

This paper provides an overview of the effects of Geometric Twist on Aircraft performance by introducing a washout condition in an aircraft wing. This condition results in an effective Angle of Attack at the wingtip that is lower than the Angle of Attack at the wing root. Using CFD analysis, the variation of aircraft performance factors such as lift, drag coefficients, and aerodynamic efficiency is calculated for different Twist angles. A plot of lift and drag coefficients at varying Angle of Attack angles has been generated based on these analyses. These results illustrate the advantages of wing twist variation, particularly at higher angles of attack. One significant advantage of Geometric Twist is that it causes stall conditions to first occur at the wing root, providing a signal to the pilot to control the aircraft before the stall reaches the wingtip. This ensures the effectiveness of control surfaces, such as Ailerons and Flaps, located at the wing trailing edge. A comparison is made between the lift, drag coefficients, and aerodynamic efficiency of twisted wings and untwisted wings with identical parameters. While twisting the wingtip yields favourable results at higher Angles of Attack (AOA) compared to an untwisted wing, the aerodynamic efficiency of the wings decreases at lower AOA. However, applying the twist angle at high angles of attack, such as 10, 12.5, 15, and 20 degrees, leads to an increase in aerodynamic efficiency.

Coupled Dynamics of Steady Jet Flow Control for Flexible Membrane Wings

We present a steady jet-flow-based flow control of flexible membrane wings for the adaptive and efficient motion of bat-inspired drones in complex flight environments. A body-fitted variational computational aeroelastic framework is adopted for the modeling of fluid–structure interactions. High-momentum jet flows are injected from the leading edge and transported to the wake flows to alter the aerodynamic performance and the membrane vibration. The coupled dynamic effect of active jet flow control on membrane performance is systematically explored. While the results indicate that the current active flow control strategy performs well at low angles of attack, its effectiveness degrades at high angles of attack with large flow separation. To understand the coupling mechanism, the variations of the vortex patterns are examined by the proper orthogonal decomposition modes, and the fluid transport process is studied by the Lagrangian coherent structures. Two scaling relations that quantitatively connect the membrane deformation with the aerodynamic loads presented in our previous work are verified even when active jet flow control is applied. A unifying feedback loop that reveals the fluid–membrane coupling mechanism is proposed. These findings can facilitate the development of next-generation bio-inspired drones that incorporate smart sensing and intelligent control.

Aerodynamic Performance of a Finite-Span Wing in a Time-Varying Freestream

The current study analyzes the dynamic response of a NACA 0015 finite-span wing to a time-varying freestream in an unsteady wind tunnel at a mean Reynolds number of 1.0×105. The aerodynamic response of the wing can be categorized based on the angle of attack and the degree of flow separation on the suction side of the wing. When the suction side is dominated by separated flow, either at low angles of attack with long laminar separation bubbles or at high, poststall, angles of attack, the lift is enhanced by the favorable pressure gradient during freestream acceleration and decremented by the adverse pressure gradient during freestream deceleration. In these cases, the sectional lift coefficient is shown to oscillate by as much as ±0.1 due to the unsteady pressure gradient. Comparatively, regions of attached flow on the wing surface remain largely unaffected by the time-varying conditions, regardless of the reduced frequency. Discrepancies between the measured lift and the estimated response from Isaacs’s theory demonstrate that the coupling between the viscous flow separation and the unsteady conditions dominates the unsteady aerodynamic response.

Effect of protuberances on the aerodynamic performance of a wind turbine blade – a review

ABSTRACT This paper thoroughly reviews the influence of protuberances on the performance of a wind-turbine blade in terms of aerodynamic characteristics, power generation capability, and noise emission. It is observed that the protuberances are significantly useful in improving the aerodynamic performance of wind turbine blades, especially in the post-stall regime and during fluctuating wind conditions, such as wind gusts. The protuberance amplitude was found to be the dominant parameter to influence the blade performance. The amplitude and wavelength were tested in the range of 2–12%c and 2–50%c, respectively. The highest wind turbine performance was observed at small amplitude and large wavelength in the pre-stall regime and at large amplitude and small wavelength in the post-stall regime. Incorporating leading edge protuberances may increase lift and power coefficients by 50 and 55%, respectively, and reduce noise emission by nearly 16 dB. With the increase in the angle of attack, the instability noise was found to change its form. The broadband noise forms at a lower angle of attack (~2°) which converts to the tonal noise at the angle of attack, 6.5°. The protuberances substantially alter the boundary layer flow, subsequently modifying the pressure and velocity field over the suction surface of the blade. A pair of vortices forms across a protuberance, which exchanges momentum with the boundary layer flow and helps delay/eliminate the flow separation. This flow separation control is responsible for the benefits as mentioned above.

On the dynamics of the turbulent flow past a three-element wing

A comprehensive analysis of the unsteady flow dynamics past the 30P30N three-element high lift wing is performed by means of large eddy simulations at different angles of attack (α = 5°, 9°, and 23°) and at a Reynolds number of Rec=750 000 (based on the nested chord). Results are compared with experimental and numerical investigations, showing a quantitatively good agreement and, thus, proving the reliability and accuracy of the present simulations. Within the slat and main coves, large recirculation bubbles are bounded by shear layers, where the onset of turbulence is triggered by Kelvin–Helmholtz instabilities. In the energy spectrum of the velocity fluctuations, the footprint of these instabilities is detected as a broadband peak; its frequency being moved toward lower values as the angle of attack increases. Kelvin–Helmholtz vortices roll-up and break down into small scales that eventually impinge into the slat and main coves lower surfaces. The slat impingement shows to be more prominent, and hence, larger velocity and pressure fluctuations are observed. The impingement strength diminishes with the angle of attack in both coves, while higher fluctuations are originated on the slat and main respective suction sides, leading to larger boundary layers. This is associated with the displacement of the stagnation point with the angle of attack. Another salient feature observed is the laminar-to-turbulent flow transition in the main and flap leading edges although the average location of this transition seems to not be affected by the angle of attack. Tollmien–Schlichting instabilities precede this transition, with the disturbances amplified by the inviscid mode at low angles of attack, while at α=23°, the local Reynolds number on the main suction side is incremented and the viscous mode becomes important. The analysis shows that the turbulent wake formed at the trailing edge of all elements dominates the dynamics downstream. This is especially true at the higher angle of attack, where a large region of velocity deficit above the flap is observed, thus indicating the onset of stall conditions.

Wind Tunnel Balance Measurements of Bioinspired Tails for a Fixed Wing MAV

Bird tails play a significant role in aerodynamics and stability during flight. This paper investigates the use of bioinspired horizontal stabilizers for Micro Air Vehicles (MAVs) with Zimmerman wing-body geometry. Five configurations of bioinspired horizontal stabilizers are presented. Then, 3-component external balance force measurements of each horizontal stabilizer are performed in the wind tunnel. The Squared-Fan-Shaped Horizontal Stabilizer (HSF-tail) is selected as the optimal horizontal stabilizer that provides the highest aerodynamic efficiency during cruise flight while maintaining high longitudinal stability on the vehicle. The integration of the HSF-tail increases the aerodynamic efficiency by more than 6% up to a maximum of 17% compared to the other alternatives while maintaining the lowest aerodynamic drag value during the cruise phase. Furthermore, balance measurements to analyze the influence of the HSF-tail deflection on the aerodynamic coefficients are conducted, resulting in increased lift force and reduced aerodynamic drag with negative tail deflections. Lastly, the experimental data is validated with CFD-RANS steady simulations for low angles of attack, obtaining a relative difference on the measurement around 5% for the aerodynamic drag coefficient and around 10% for the lift coefficient during the cruise flight that demonstrates a high degree of accuracy in the aerodynamic coefficients obtained by external balance in the wind tunnel. This work represents a novel approach through the implementation of a horizontal stabilizer inspired by the structure of the tails of birds that is expected to yield significant advancements in both stability and aerodynamic efficiency, with the potential to revolutionize MAV technology.

Experimental and numerical investigation on the unsteady interaction in longitudinal stage separation for parallel-staged two-body configuration

Hypersonic stage separation is a significant process, probably involving complex aerodynamic interaction, which determines the survival of two-stage-to-orbit (TSTO) vehicles. The longitudinal stage separation (LSS) scheme is proposed to improve the safety of the parallel-staged TSTO vehicle, where a small interstage gap may result in weak interference. Therefore, an experimental and numerical study of LSS for the parallel-staged TSTO vehicle at Mach 7 with different angles of attack (AoA) is carried out. The dynamic interaction, including variations in the shock structure, wall pressure distribution, and unsteady aerodynamics, is investigated by testing and numerical simulation. The LSS experiments for the TSTO vehicle were performed using a high-speed pneumatic ejection launch system in the JF-12 shock tunnel, and the method was developed using high-speed visualization and image processing techniques to capture the separating trajectory. The numerical simulations were carried out using the overset grid method and solving the Navier–Stokes equations coupled with the rigid body dynamics equations to obtain the laminar flows over the TSTO vehicle during LSS. The qualitative and quantitative comparison of the test and numerical results showed good agreement in terms of aerodynamic performance, flowfield pattern, wall pressure, and separation trajectory. They show that the small interstage gap of the LSS leads to weak type I and VI shock–shock interactions, with short-duration weak shock reflection at a higher AoA. Furthermore, no shock reflection or interstage gap is observed at lower AoA. Moreover, no stage recontact is observed, and the safety and feasibility of LSS for parallel-staged TSTO vehicles are demonstrated.

Research on transition corridor of distributed tilting power configuration

For a distributed tilt-power configuration aircraft, the transition corridor was studied based on limited conditions such as wing lift characteristics and propeller power requirement. By establishing the dynamic model of each component and balancing according to the balance equation, the boundary of the transition corridor of distributed tilting dynamic configuration is determined. The results show that the minimum forward flight speed required for a low Angle of attack is less than that for a high Angle of attack; the wing lift characteristic high-speed boundary and the propeller power limit boundary together form the high-speed boundary of the transition corridor.

Investigation of the aerodynamic effects of bio-inspired modifications on airfoil at low Reynolds number

A numerical study was performed to investigate flow behaviors around bio-inspired modified airfoils compared with NACA 4412 airfoil at Re=5.8x104 by solving the two-dimensional, RANS equations with k-ω STT turbulence model. The obtained results reveal a rather abrupt decrease of lift at stall for the NACA 4412 airfoil in contrast to the mild stall depicted by the top-modified airfoil. As compared to the experimental results of the profiled airfoil in the literature, the characteristic behavior of the variation in the lift coefficient shows resemblance. It is seen that from the velocity distribution results, fluid flowed smoothly along the streamlined nose of NACA 4412 airfoil until α=4⁰ and streamlines adhered well for both airfoils at low angles (0⁰, 2⁰). Smaller circulation bubbles were noticed to settle in the canyons of the corrugated cross-section of the top-modified airfoil. In the wake region of the modified airfoil, there is no obvious large flow separation or circulation region at low angles of attack. However, the blue regions of the dimensionless velocity over the NACA 4412 airfoil and bottom-modified airfoil were narrower than over the top-modified airfoil. The recirculation zone over the airfoil started to enlarge, and the rolling up of the trailing-edge vortex appeared. After α=12⁰, the adverse pressure gradient on the suction side of the airfoils became more intense. In the wake zones, it was seen that the circulation regions grew remarkably and became largest as the angle of attack rose to α=16⁰, which pointed out increased drag forces of airfoils.

Numerical Study on Combustion-Driven Jet Actuation for Aerodynamic Control of Airfoil Flows

In this study, a numerical investigation is conducted on combustion-driven pulsed-jet actuation to control the flow around a lifting surface. Based on relevant experimental measurements and computations, high-speed jets are generated from the impulsive variation in pressure at the actuator boundary. A supersonic jet flow is momentarily generated by combustion in a reaction chamber of the actuator, and the flow interacts with the external flow around the lifting surface and alters the aerodynamic characteristics. The computational results indicate that the flow control performance of the jet actuation is significant at a high-incidence angle of attack, such as beyond the stall angle, whereas the impact is minimal at low angles of attack, such as in the linear lift region. Repetitive jet actuation can produce additional momentum to the external flow and alters the pressure distribution on the suction surface, particularly downstream of the actuator location. This pressure variation from the actuation yields an additional lift force on the lifting surface and reduces the amplitude of the aerodynamic moment at a given angle of attack, thus enhancing the aerodynamic performance of the airfoil.