Unmanned aerial vehicles (UAVs) with flying wing designs offer unique advantages in terms of maneuverability and fuel efficiency, but achieving precise control over these tailless aircraft can be challenging. This study investigates the potential of notch leading edge (NLG) technology to improve the effectiveness of control surfaces in flying wing UAVs. NLG refers to a design feature where notches or depressions are strategically incorporated along the leading edge of the wings. This seemingly minor modification can significantly impact the aircraft's aerodynamic performance and controllability. The present research leverages numerical methods to meticulously examine the influence of NLG placement on a crucial parameter: the roll torque coefficient. This coefficient quantifies the effectiveness of control surfaces, such as split drag rudder, in yawing moments, essential for maneuvering the aircraft. The investigation adopts a parametric approach, meticulously varying the dimensions and positions of the NLG along the wing span. This comprehensive analysis reveals a fascinating relationship between NLG placement and control surface efficacy. The study convincingly demonstrates that positioning the NLG closer to the wing root, where the wing meets the fuselage, yields superior results compared to placements near the wingtip.

Axisymmetric flows present unique numerical challenges absent in planar configurations, with bow shocks positioned closer to bodies, compressing stabilizing subsonic layers and promoting carbuncle instabilities. This study conducts the first systematic evaluation of three AUSM-family schemes (AUSM+UP, SLAU, AUSM+M) across inviscid and viscous axisymmetric flows spanning transonic to supersonic regimes. Four benchmarks are examined: supersonic sphere (M=4), convergent-divergent nozzle, mixed-compression inlet (M=2) and transonic bump (M=0.875). AUSM+M emerges superior for high-speed axisymmetric applications, uniquely suppressing carbuncle instability on the sphere while achieving 65.8% improvement in post-shock Mach accuracy and 12% shock stand-off error vs AUSM+UP's 28.5%. While AUSM+M incurs a higher computational cost in simpler flows (requiring 54% more CPU time for the sphere case), this expense is justified by its superior capability to suppress carbuncle instabilities and ensure numerical robustness, a trade-off that becomes advantageous in complex configurations where it achieves faster convergence. For the complex inlet, AUSM+M converges 21.3% faster with optimal total-pressure recovery (3.7% error). It captures shock-induced separation most accurately, predicting separation length within 2.75% on the bump. Conversely, AUSM+UP exhibits severe shock instabilities, while SLAU consistently fails to resolve strong shocks despite excellent low-Mach performance. Results establish AUSM+M as the preferred scheme for axisymmetric simulations where shock resolution and numerical stability are paramount.

This paper introduces NAUSM+M+AUFS (New Improved Advection Upstream Splitting Method Plus Artificially Upstream Flux Vector Splitting), a novel hybrid computational scheme for simulating compressible flows on triangular grids. The AUSM+M (Improved Advection Upstream Splitting Method) method is enhanced through two key modifications to boost numerical stability and robustness in high Mach number and hypersonic flows. The first modification redefines the interfacial numerical sound velocity, reducing shock anomalies and improving shock-capturing by integrating velocity and characteristic sound speed parameters. The second modification addresses the insufficiency of the pressure flux dissipation term at supersonic speeds by introducing a formulation that increases dissipation proportionally to the Mach number, thereby enhancing performance in high-speed flows. These enhancements constitute the NAUSM+M method. The NAUSM+M+AUFS scheme combines the strengths of NAUSM+M and AUFS (Artificially Upstream Flux Vector Splitting) methods, particularly in overcoming the limitations of NAUSM+M in handling shock instabilities and the carbuncle phenomenon on structure triangular grids. A dynamic switching function adjusts the weighting between NAUSM+M and AUFS, optimizing accuracy and stability based on local flow conditions. Numerical tests demonstrate that NAUSM+M+AUFS significantly outperforms AUSM+M, NAUSM+M, and AUFS, effectively eliminating the carbuncle phenomenon and providing smooth shock wave contours. In steady flow analysis, the new hybrid method achieves convergence speeds comparable to AUFS and shows 15% to 45% superior convergence accelerating than AUSM+M, depending on the convergence rate. In addition, in steady flow analysis, the accuracy of NAUSM+M+AUFS is 46% better than that of AUFS. This approach represents a significant advancement, offering a robust, accurate, and efficient solution for high-speed aerodynamic simulations, with broad applicability across various compressible flow challenges.

Wing smarting and eliminating conventional control surfaces are fundamental parameters for improving aircraft aerodynamic performance in future aviation. Twist is a well-known tool that, along with the development of morphing technology, can play a crucial role in controlling next-generation aircraft. However, wing twisting with a control approach requires many aerodynamic studies, particularly at the high aft-swept angle; this is more noticeable in the flying wing configuration. In this paper, a control strategy based on twist morphing has been evaluated aerodynamically at the Swing, a flying wing configuration. Extraction of aerodynamic coefficients and flow field on the wing have been performed using the computational fluid dynamics (CFD) method in an incompressible flight regime. To comprehensively cover the control needs, two control concepts have been introduced, called independent and nonindependent twist. Within the concept of independent twist, a co-oriented twist arrangement (Co-OTA) and counteroriented twist arrangement (Cun-OTA) are applied to the wings, which are used to produce pitching and rolling moments, respectively. The results show these control arrangements have high efficiency at low angles of attack (AoAs), but as the AOA increases, their aerodynamic performance will gradually decrease. In this regard, a significant challenge for Cun-OTA is producing the yawing moment during roll maneuvering. Attempts to solve this problem have led to the idea of a compound twist arrangement (CTA). The existence of a control arrangement to generate yawing moment is beyond the capacity of the independent twist concept. The nonindependent twist has been introduced as a solution to this issue. Conventional control surfaces have various weaknesses, one of the most important of which is the unfavorable effects on aerodynamic efficiency; the reduction of aerodynamic efficiency leads to a drop in flight endurance and an increase in fuel consumption. At a step forward, with the approach of wing smarting and morphing technology, the tasks of conventional control surfaces can be delegated to the wing itself. Therefore, it will be possible to remove the control surfaces and integrate the wing. In this article, according to the mentioned approach, an attempt was made to introduce a control strategy based on the geometric twist. In this strategy, the value and direction of the twist applied to each wing produces a specific control arrangement. Control arrangements (three types of arrangements) are involved in producing longitudinal and lateral and modified lateral moments. The results indicate that the use of twisting in aircraft control is efficient for a wide range of flight conditions, and this capability exists to be used in the next generation of aircraft. In this way of control, challenges are deliberately related to the design of the operating mechanism, safety issues, and weight.

This paper introduces a novel hybrid numerical method, SAUSM, designed for accurate and robust simulation of compressible flows governed by the Euler equations. While the AUSM[Formula: see text] scheme provides proper resolution of smooth flow features, it is susceptible to anomalies, particularly the carbuncle phenomenon near strong shock discontinuities. Conversely, the AUFS scheme offers inherent stability in capturing shocks; however, it lacks the accuracy of AUSM[Formula: see text] in smooth regions. The proposed SAUSM method combines AUSM[Formula: see text] and AUFS through an adaptive weighting function, facilitating a seamless transition between the schemes. This approach preserves the accuracy of AUSM[Formula: see text] in smooth regions while ensuring robust shock-capturing capabilities near discontinuities. The effectiveness of the SAUSM method is rigorously demonstrated through a comprehensive suite of progressively complex test cases. Numerical experiments demonstrate SAUSM’s proficiency in resolving intense shock patterns and discontinuities without introducing anomalies. In the selected test cases, SAUSM agrees with reference solutions and effectively mitigates anomalies observed in AUSM[Formula: see text], including kinked Mach stems. In the challenging test case involving hypersonic blunt body flow over a cylinder, SAUSM adapts dissipation effectively by utilizing its adaptive weighting function to generate smooth pressure distributions, thereby eliminating the carbuncle instability linked to AUSM[Formula: see text] when applied to a high aspect ratio grid. The consistent formulation of flux splitting and the adaptive weighting in SAUSM prevent excessive dissipation away from discontinuities, thus preserving accuracy comparable to that of exact Riemann solvers. Consequently, SAUSM emerges as a promising and innovative approach to accurately and robustly simulate a wide range of compressible Euler flows. The comprehensive results obtained from the validation tests firmly establish SAUSM as a highly effective general-purpose technique for computational fluid dynamics in academic research.

Three preconditioning methods suggested by Eriksson, Choi, Merkel, and Turkel have been utilized within a 2D upwind Euler flow solver designed for unstructured grids. These strategies efficiently address the complexities of steady inviscid flows at low Mach numbers. The conservative formulations of the preconditioning matrices are rigorously derived. This implementation enables a more accurate evaluation of high-gradient flows. Extensive simulations are conducted on various flow scenarios, including flows over the NACA0012 airfoil, a multi-element three-element airfoil, and a smooth bump with varying Mach numbers, to validate the effectiveness of the aforementioned preconditioning strategies. Compared to the non-preconditioned approach, the results demonstrate significant accuracy and convergence speed improvements for all three preconditioning methods. These strategies exhibit remarkable efficiency for low Mach and incompressible flows. Among the three approaches, the Turkel preconditioner stands out with its optimal condition number, leading to superior performance. For low Mach numbers, convergence is accelerated by up to 88%, while at transonic speeds, it still achieves a notable 38% increase in convergence speed. Additionally, the preconditioning techniques preserve solution accuracy near challenging stagnation points. This study establishes a unified conservative framework for assessing preconditioning approaches and highlights their ability to resolve the complex fluid physics of low Mach number flows on unstructured grids. The findings underscore the significance of employing such strategies to enhance accuracy and computational efficiency in evaluating high-gradient flows.

This study presents an innovative implicit–explicit time-stepping algorithm based on a first-order temporal accuracy method, addressing challenges in simulating all-regimes of fluid flows. The algorithm's primary focus is on mitigating stiffness inherent in the density-based “Roe” method, pivotal in finite volume approaches employing unstructured meshes. The objective is to comprehensively evaluate the method's efficiency and robustness, contrasting it with the explicit fourth-order Runge–Kutta method. This evaluation encompasses simulations across a broad spectrum of Mach numbers, including scenarios of incompressible and compressible flow. The scenarios investigated include the Sod Riemann problem to simulate compressible Euler equations, revealing the algorithm's versatility, and the low Mach number Riemann problem to analyze system stiffness in incompressible flow. Additionally, Navier–Stokes equations are employed to study viscous and unsteady flow patterns around stationary cylinders. The study scrutinizes two time-stepping algorithms, emphasizing accuracy, stability, and computational efficiency. The results demonstrate the implicit–explicit Runge–Kutta algorithm's superior accuracy in predicting flow discontinuities in compressible flow. This advantage arises from the semi-implicit nature of the equations, reducing numerical errors. The algorithm significantly enhances accuracy and stability for low Mach number Riemann problems, addressing increasing stiffness as Mach numbers decrease. Notably, the algorithm optimizes computational efficiency for both low Mach number Riemann problems and viscous flows around cylinders, reducing computational costs by 38%–68%. The investigation extends to a two dimensional hypersonic inviscid flow over cylinder and double Mach reflection case, showcasing the method's proficiency in capturing complex and hypersonic flow behavior. Overall, this research advances the understanding of time discretization techniques in computational fluid dynamics, offering an effective approach for handling a wide range of Mach numbers while improving accuracy and efficiency.

In the present study, the effects of the wing fence on the wing tip vortices and control surfaces located at the tip of the wing in a flying wing aircraft have been investigated using a numerical method. For the size of the fences, the average dimensions extracted from the wing tip vortices at different angles of attack are used. The basic determining parameter is the rolling torque coefficient, which is tried to be shown by employing a parametric study of the flow behavior in different situations of fence placement. These effects on the rolling torque of the aircraft are measured due to the presence of the split drag rudder control system. In this study, the fences were installed at three different heights and three different positions along the length of the wing, which were investigated at angles of attack of 7 to 16 degrees. The next stage of the research is to design the dimensions of the fence using the single-objective optimization method (a method to find the best solution for a problem with a specific goal). The designing of the fences at three points based on the dimensions of the wing tip vortex is carried out with the computational fluid dynamics (CFD) method (CFD is a computational method that uses physical laws to predict the behavior of fluids.). The aim of this research is to achieve the best design that converges to an optimal solution with minimum time and cost (CFD solution is long). However, CFD analysis requires a lot of computational time. To address this challenge, we employed a hybrid learning model comprising the radial basis function (RBF), a type of artificial neural network, and Kriging, a Gaussian process-based interpolation technique. The dataset for training the hybrid model was obtained from numerical solutions of CFD simulations involving a fence placed at various locations on the wing. Additionally, a genetic algorithm was employed as the optimization method in all instances where it was required. Using the power of machine learning techniques helped us identify the optimal placement of the fence to prevent it from being engulfed by the vortex and to optimize the utilization of the split drag system, yielding significant improvements.

In this study, models for the wake capture lift and drag force coefficients of a hovering flapping wing were presented using numerical fluid dynamics simulation to improve the blade element theory. The investigated wing is inspired by the fruit fly and has combined flapping and pitching movements. The effect of changing the wing acceleration time at the start ([Formula: see text]) and end of the half cycle ([Formula: see text]), as well as the Reynolds numbers in the range of 136–6800, on wake capture using the Taguchi orthogonal array test design, is investigated using the numerical fluid dynamics method, and the values of average lift and drag coefficients due to wake capture are obtained. These force coefficients were applied to linear and nonlinear regression methods to obtain the mathematical model, and a model for its changes was extracted. Finally, to obtain the instantaneous coefficients, the extracted models were placed in the normal distribution function, and the final instantaneous model was obtained. Examining the verification cases of the application of these wake capture relationships with the blade element theory, as well as the effects of translational force, rotational force and added mass force, revealed that this developed theory is capable of correctly predicting the wake capture force’s initial peak. The force coefficient trend in the final quasi-steady model with wake capture is similar to the computational fluid dynamics (CFD) results, according to a qualitative examination of the lift and drag force coefficients in a half cycle. This demonstrates a significant result: this theory, which is divided into four parts: transnational, rotational, added mass and wake capture, adequately covers the general physics of these complex movements.

This paper introduces a novel platform that integrates three preconditioning matrices based on conservative variables: the Turkel, Choi–Merkle and Onur matrices. The platform aims to compare these matrices in terms of accuracy and robustness by investigating their performance in solving three distinct and challenging high-gradient laminar flow problems: (i) Bi-Plane NACA0012 airfoil, (ii) lid-driven flow in a square cavity and (iii) flow in a planar T-junction. These problems serve as new and challenging test cases to accurately determine the abilities of the preconditioning matrices. The preconditioning matrices are applied to evaluate the numerical solutions, and their performance in complex flow fields is assessed in terms of accuracy and efficiency. By solving these flow problems, the effectiveness of the preconditioning matrices is thoroughly analyzed. By integrating these preconditioning matrices into a single platform, this paper significantly contributes to the field. The approach enables a direct and meaningful comparison of the performance of the Turkel, Choi–Merkle and Onur matrices in solving these new and challenging laminar flow problems. While all three matrices demonstrate comparable accuracy in predicting flow characteristics like pressure coefficients, Turkel’s method shows superior convergence acceleration across the almost test cases due to alpha parameter and modifications of the momentum equations. In the external flow case, Turkel converges 22–53% faster than the other matrices by adjusting momentum terms with an alpha parameter. For the internal lid-driven cavity case, Turkel again accelerates convergence up to 38% over Choi–Merkle as Reynolds and Mach numbers increase. However, Onur’s method stalls at high Reynolds/Mach numbers. At low values of Reynolds and Mach numbers, Onur reduces computational cost by 17%. Finally, for the T-junction case, Turkel and Choi–Merkle perform almost identically, decreasing CPU time by 55% vs Onur, which lacks convergence robustness. While the preconditioning matrices have similar accuracy, Turkel offers the best convergence improvement by accounting for momentum effects. Onur works well at low Reynolds/Mach numbers but shows limitations at higher values. The selection of the optimal preconditioning matrix should be based on whether momentum or viscosity dominates in the flow field, as well as the intensity of the viscosity gradient rate.

Purpose Ground effect is one of the important factors in the enhancement of wing aerodynamic performance. This study aims to investigate the aerodynamic forces and performance of a flapping wing with the bending deflection angel under the ground effect. Design/methodology/approach In this study, the wing and flapping mechanism were designed and manufactured based on the seagull flight and then assembled. It is worth noting that this mechanism is capable of wing bending in the upstroke flight as big birds. Finally, the model was examined at bending deflection angles of 0° and 107° and different distances from the surface, flapping frequencies and velocities in forward flight in a wind tunnel. Findings The results revealed that the aerodynamic performance of flapping wings in forward flight improved due to the ground effect. The effect of the bending deflection mechanism on lift generation was escalated when the flapping wing was close to the surface, where the maximum power loading occurred. Practical implications Flapping wings have many different applications, such as maintenance, traffic control, pollution monitoring, meteorology and high-risk operations. Unlike fixed-wing micro aerial vehicles, flapping wings are capable of operating in very-low Reynolds-number flow regimes. On the other hand, ground effect poses positive impacts on the provision of aerodynamic forces in the take-off process. Originality/value Bending deflection in the flapping motion and ground effect are two influential factors in the enhancement of the aerodynamic performance of flapping wings. The combined effects of these two factors have not been studied yet, which is addressed in this study.

The study introduces an innovative approach that combines dynamic and thermal lattice Boltzmann simulations utilizing the ghost fluid boundary detection method for enhanced flow stability during flapping oscillation. This novel methodology is applied to large-eddy simulations of a flapping airfoil, aiming to capture the complex dynamics of oscillatory vortices and their impact on aerodynamics, while also allowing control over aerodynamic responses through airfoil surface temperature modulation. The investigation involves integrating the ghost fluid method into the lattice Boltzmann framework, synchronized with each time step to accurately model both the airfoil's oscillatory and boundary-induced movements. The advancement of specific fluid domain nodes due to boundary motion is managed using a recharging and force imposition scheme, estimating energy, and density function distributions on these nodes. Variations in lift and drag coefficients resulting from dynamic viscosity changes and thermal effects are observed due to airfoil surface temperature adjustments. A nonreflecting boundary condition is introduced to regulate flow velocity upper and lower near the domain boundary, facilitating smooth flow transfer from the boundary to the oscillating airfoil flow and minimizing transverse interference. Changes in energy exchange lead to delayed boundary layer separation, suggesting enhanced performance with reduced airfoil temperature. It is shown that decreasing the airfoil temperature by 100 K compared to the fluid temperature reduces the drag coefficient by 34% and increases the lift coefficient by 14%, while with an increase in 100°, the drag coefficient increases by 14.9% and the lift coefficient decreases by 4%. The proposed approach offers computational simplicity, concise solvable equations, and high accuracy, eliminating the need for mesh size adjustments when simulating different Reynolds numbers. Additionally, its ability to accommodate heat transfer-induced alterations within the aerodynamic context is highlighted. Comparative analysis with the finite volume method validates its effectiveness, demonstrating potential applications for controlling aerodynamic coefficients through controlled thermal interventions. In conclusion, the study presents a comprehensive methodology that integrates dynamic and thermal lattice Boltzmann simulations with the ghost fluid boundary detection method for enhanced flow stability during flapping oscillation. The insights gained contribute to a deeper understanding of complex aerodynamic phenomena, with implications for aerospace and fluid dynamics research, where accurate prediction and control of airfoil behavior are crucial. In summary, this study offers a groundbreaking strategy that seamlessly integrates dynamic and thermal lattice Boltzmann simulations, leveraging the ghost fluid boundary detection method for enhanced stability in flapping oscillatory motion. The outcomes contribute to a deeper understanding of intricate aerodynamic phenomena, thus holding promise for broader applications in aerospace and fluid dynamics research.

Purpose This study aims to investigate the effectiveness of two types of winglets, multi-tip and raked, on the performance of sinusoidal and simple leading-edge wings and compares it by a numerical method. Design/methodology/approach The wing configuration in this study is rectangular and uses NACA0020 section, and all simulations are performed by a numerical method based on finite volume and base pressure algorithm in Reynolds 2 × [10]^5. In the mentioned numerical method, the flow is considered turbulent, and the k-ω-SST model is used. To calculate the stresses on the wing surface, the mesh is extended to below the viscous layer, and a second-order upstream accuracy is used to calculate the convection flux. Findings The use of raked and multi-tip winglets for the sinusoidal edge of the wing improved aerodynamic performance by 5.12 and 2.28%, respectively, and the greatest effect of these two winglets was on increasing the lifting force and reducing the inductive drag, respectively. Also, by examining the distribution of induced vortices around the configurations, it was found that the curvature of the sinusoidal wing tip at the angles of attack before stall reduced the strength of the induced vortices and, the use of winglet during and after stall, caused increased aerodynamic performance of the sinusoidal wing. Practical implications The whale is an international species of aquatic animal found in most of the world’s oceans. It has large fin aspect ratios that have a series of bulges at the edge of the attack, which improves the aerodynamic performance near and after stall. Today, one of the fields of research is the use of this idea in the wings of micro air vehicle. Originality/value Winglet reduces induced drag in simple wings. So far, the effect of winglets on wings with sinusoidal attack edges has not been investigated.

The vital role of the numerical scheme is becoming increasingly critical as the use of computational fluid dynamics grows. To address the unfavorable effects experienced in low-speed flows when using the AUSM+M scheme (Improved Advection Upstream Splitting Method), the present paper presents an improved approach known as Modified-AUSM+M (M-AUSM+M). This novel method offers enhanced reliability in simulating low-Mach number flows, effectively mitigating the challenges associated with low-speed symptoms encountered in the original AUSM+M scheme. The novel scheme is facilitated by the parameter-free form of the pressure diffusion term in the mass flux and the low-dissipative form of the velocity diffusion term in the pressure flux. The impacts of these critical ingredients are then thoroughly evaluated, and the different characteristics are explored in terms of robustness and accuracy using a wide range of low-Mach test cases. The proposed scheme maintains a consistent correlation between accuracy and convergence speed. In addition, the recently devised technique demonstrates superior accuracy compared to AUSM+M and AUSM+UP schemes when dealing with low-Mach flows. Furthermore, the findings indicate an incredible reduction in iteration numbers, ranging from 30% to 80%, by employing the enhanced scheme in low-Mach domains. In the investigation of high-Mach test cases, the newly developed method preserves the accuracy achieved by AUSM+M in high-Mach flows.

The geometric structure of birds' wings, especially edges and wingtips, plays an important role in improving flight conditions. In order to investigate the changes in flow physics caused by the corrugated trailing edge and wingtip slots, a three-dimensional numerical simulation of the eagle wing with corrugated trailing edge and flapping motion was carried out in forward flight. First, lift and drag forces were compared between numerical and experimental results, showing acceptable agreement. As key parameters, the lift force, thrust force, and lift-to-drag ratio were compared between the modified and the baseline wings. The results revealed that using the modified wing led to the improvement of the lift force by 14% on average compared to the baseline wing; with a maximum increase of 45%. Furthermore, the thrust force of the modified eagle wing was higher than that of the baseline wing by up to 21%. According to the analysis of the flow physics in the present simulation, the pressure difference between the upper and lower surface of the modified eagle wing was larger compared to the baseline wing, and the highest effect on the flow physics was exerted by wingtip slots. Also, the behavior of vortices at different sections across the wing length revealed that the interaction between the wingtip slots led to flow deviation in this region, an increase in the flow kinetic energy in the boundary layer, and a decrease in the flow separation and vortices separated from the wingtip. Further, it caused the separated vortices from the wingtip to change from the concentrated state to smaller vortices with less strength, which reduced the induced drag and enhanced aerodynamic coefficients.

The aerodynamic performance of fabricated eagle wing with the corrugated trailing edge was investigated experimentally and numerically in this study. In this respect, wings were designed by imitating the bionic eagle wing and fabricated using a 3D printer. Tests on the wings were performed in a wind tunnel for different bending deflection angles, flapping frequencies, angles of attack, and forward flight velocities. As the key parameters, lift and thrust forces, lift-to-drag and input power were compared between eagle wing and simplified wing. The results revealed that the eagle wing with the corrugated trailing edge led the lift force to be improved by 14% on average compared to the simplified wing, with the largest improvement of 45%. Further, the thrust force of a eagle wing was greater than that of the simplified wing (up to 21%). Also, the input power of the simplified wing mechanism was greater than that of the eagle wing. Numerical simulations were performed to study flow physics. After validating the numerical results against the experimental results, the lift and thrust force, lift-to-drag, pressure contours and vortex were compared between the eagle and simplified wings. The results of the numerical simulation revealed that lift-to-drag increased with for eagle wing with the corrugated trailing edge of eagle wing. the pressure difference between the upper and lower surfaces of the eagle wing was larger than that of the simplified wing; therefore, the lift force produced by the eagle wing was larger than that of the simplified wing.

In this investigation, finite difference lattice Boltzmann method (FD_LBM) is developed to solve heat transfer effect behavior on the symmetrical and unsymmetrical airfoils with plunge oscillations.In this simulation, the equations of motion and energy are executed using LBM and FD simultaneously.The LB method is integrated with ghost flow for predicted curve boundary.The ghost flow method is a Cartesian-based method that, in addition to being practical and straightforward, retains many advantageous features of structured meshes, can be used for complex geometries, and has a high degree of flexibility.In other words, when the body oscillates, it is important to determine its position caused by the change in the mesh structure at any time.While the ghost method detects the object's position well, the new technique can capture the details of flow more accurately and stably than the other methods.Combining the ghost method with LBM provides a new technique that can investigate thermal behavior's effect on the airfoil with greater accuracy and stability.This combination of modern methods with high accuracy and stability in complex geometries has not been studied.The results are compared with the literature and show that this method has better convergence in different Reynolds and temperatures with changes at boundary conditions in the airfoil.

Since laminar separation bubbles are neutrally shaped on the suction side of full-span wings in low Reynolds number flows, a roughness element can be used to improve the performance of micro aerial vehicles. The purpose of this article was to investigate the leading-edge roughness element’s effect and its location on upstream of the laminar separation bubble from phase portrait point of view. Therefore, passive control might have an acoustic side effect, especially when the bubble might burst and increase noise. Consequently, the effect of the leading-edge roughness element features on the bubble’s behavior is considered on the acoustic pressure field and the vortices behind the NASA-LS0417 cross-section. The consequences express that the distribution of roughness in the appropriate dimensions and location could contribute to increasing the performance of the airfoil and the interaction of vortices produced by roughness elements with shear layers on the suction side has increased the sound frequency in the relevant sound pressure level (SPL). The results have demonstrated that vortex shedding frequency was increased in the presence of roughness compared to the smooth airfoil. Also, more complexity of the phase portrait circuits was found, retrieved from velocity gradient limitation. Likewise, the highest SPL is related to the state where the separation bubble phenomenon is on the surface versus placing roughness elements on the leading edge leads to a negative amount of SPL.

Although the tubercle wings provide good maneuverability at post-stall conditions, the aerodynamic performance at pre-stall angles is threatened by forming a laminar separation bubble at the trough section of the tubercle wing; consequently, the flight endurance and range are reduced. In the present study, the idea of passive flow control is introduced by using the distribution of static roughness elements on a full-span wing with a sinusoidal leading edge. Initially, the effect of roughness element length, height, and its location are studied at a pre-stall angle (16-degree). Their effect on the laminar separation bubble and vortex shedding formed behind the wing are also investigated. The Reynolds number is assumed to be equal to [Formula: see text] which is in the range of critical Reynolds number and matches to the micro aerial vehicles application. An improved hybrid model, improved delay detached eddy simulation IDDES, has been used to model the flow turbulence structure. In the extended transition region at low Reynolds numbers, the roughness bypassed the instability. Consequently, roughening the surface of the aerofoil increased the boundary layer’s flow momentum, making it more resistible to adverse pressure gradients. By suppressing the bubble, the static roughness element led to pre-stall flow control, which saw an increase in lift coefficient, [Formula: see text], and a decrease in drag coefficient, [Formula: see text]. The results have been demonstrated that the aerodynamic performance, [Formula: see text], has been improved approximately 22.7%, 38%, and 45% for [Formula: see text], and [Formula: see text], respectively. The optimal arrangement of static roughness elements could decline the size of the vortices and strengthen the cores associated with them. This claim can be interpreted with the vortex shedding frequency.

Sinusoidal leading-edge wings have shown a high performance after the stall region. In this study, the role of smart flaps in the aerodynamics of smooth and sinusoidal leading-edge wings at low Reynolds numbers of 29,000, 40,000 and 58,000 is investigated. Four wings with NACA 63 4 -021 profile are firstly designed and then manufactured by a 3 D printer. Beam bending equation is used to determine the smart flap chord deflection. Next, wind tunnel tests are carried out to measure the lift and drag forces of proposed wings for a wide range of angles of attack, from zero to 36 degrees. Results show that using trailing-edge smart flap in sinusoidal leading-edge wing delays the stall point compared to the same wing without flap. However, a combination of smooth leading-edge wing and smart flap advances the stall. Furthermore, it is found that wings with smart flap generally have a higher lift to drag ratio due to their excellent performance in producing lift.