High Angles Of Attack Research Articles

Transonic leading-edge stall flutter: modelling, simulations and experiments

This work is a numerical and experimental study of a rectangular thin plate undergoing stall flutter at Mach 0.8. This constitutes one of the first studies of this kind where three-dimensionality is fully implemented in a numerical simulation including the test-section effects characterizing wind-tunnel experiments. In order to break down the fluid–structure interaction to its main driving phenomena, an aerodynamic model is proposed that is based on computationally inexpensive steady-state simulations. Two types of dynamic instability are observed in the numerical simulations; Flutter by mode coalescence is promoted at zero flow incidence, however, high bending precludes this from happening for higher values of angle of attack. Stall flutter is instead a nonlinear one-degree type of instability. Both of these instability mechanisms can be explained in terms of hysteretic behaviour of the pressure distribution, which becomes more pronounced at high angles of attack, when a large separation region is formed. Tests were conducted employing titanium alloy plates in order to survive the aerodynamic loads characterizing the wind-tunnel initial transient. However, due to wall interference, high bending was promoted so that the internal stress exceeded the yield values before flutter could be measured. Numerical simulations were in general agreement with the experiment in terms of both amplitude and oscillation frequency.

Identification of control equations using low-dimensional flow representations of pitching airfoil

This study investigates the application of data-driven modeling techniques for understanding the complex dynamics of pitching airfoils at low Reynolds numbers and high angles of attack. Linear and nonlinear dimensionality reduction methods, namely principal component analysis (PCA) and isometric mapping (ISOMAP), are employed to obtain low-dimensional representations of the flow field. Subsequently, sparse identification of nonlinear dynamics (SINDy) is utilized to model the governing equations. The key findings are as follows: PCA primarily captures linear information, with the first two to three dimensions maintaining relatively low reconstruction errors. In contrast, ISOMAP excels in capturing nonlinear features, exhibiting noticeably smaller reconstruction errors. The main information is concentrated in the two-dimensional plane constructed by PCA1 and PCA2 (or ISOMAP1 and ISOMAP2). Differences in trajectory planes formed by combinations of other axes reflect flow field disparities. ISOMAP provides a nonlinear low-dimensional representation, advantageous for capturing nonlinear relationships between flow field characteristics and governing equations. The combination of ISOMAP and SINDy yields virtually no errors in identifying governing equations. Conversely, PCA and SINDy result in significantly different linear trajectories, leading to higher reconstruction errors. The identified governing equations using ISOMAP and SINDy remain consistent across different datasets, demonstrating the method's stability and robustness in accurately characterizing flow field properties under similar conditions.

Effects of cornering conditions on the aerodynamic characteristics of a high-performance vehicle and its rear wing

This study investigates the aerodynamic behavior of a high-performance vehicle and the interaction with its rear wing in straight-line and steady-state cornering conditions. Analyses are performed with Reynolds-averaged Navier–Stokes based computational fluid dynamics simulations using a moving reference frame and overset mesh technique, validated against moving ground wind tunnel experiments. The results indicate a significant 20% decrease in downforce and 35% increase in drag compared to straight-line conditions at the smallest considered corner radius of 2.9 car-lengths. Downforce losses primarily stem from performance deficits on the underbody and rear wing, alongside elevated upper body lift. Drag penalties mainly result from additional pressure drag induced by a recirculation wake vortex generated behind the vehicle's inboard side. The vehicle's lateral pressure distribution is also affected, introducing a centripetal force that increases with smaller corner radii. Additionally, analyses of the rear wing reveal alternations of its aerodynamic characteristics in cornering, particularly impacting vortical flow and suction on the lower surface. Throughout the operating conditions, the rear wing's individual downforce contribution falls off beyond its stall angle. At higher angles of attack, the rear wing primarily generates downforce by pressurizing the vehicle's upper surfaces, but its interaction with the near-wake leads to a substantially increased pressure drag. Overall, these findings provide crucial insights into the intricate aerodynamic interactions of high-performance vehicles in diverse operating conditions as well as form an essential foundation for future research on static and active aerodynamic designs in the pursuit to optimize vehicle performance in dynamic driving conditions.

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.

Embedded shear layers in turbulent boundary layers of a NACA0012 airfoil at high angles of attack

An investigation of turbulent boundary layers (TBLs) is presented for a NACA0012 airfoil at angles of attack 9 and 12 deg. Wall-resolved large eddy simulations (LES) are conducted for a freestream Mach number M=0.2 and chord-based Reynolds number Re=4×105, where the boundary layers are tripped near the airfoil leading edge on the suction side. For the angles of attack analyzed, mild, moderate and strong adverse pressure gradients (APGs) develop over the airfoil. Despite the strong APGs, the mean flow remains attached along the entire airfoil suction side. Similarly to other APG-TBLs investigated in the literature, a secondary peak appears in the Reynolds stress and turbulence production profiles. This secondary peak arises in the outer layer and, for strong APGs, it may overcome the first peak typically observed in the inner layer. The analysis of the turbulence production shows that other components of the production tensor become important in the outer layer besides the shear term. For moderate and strong APGs, the mean velocity profiles depict three inflexion points, the third being unstable under inviscid stability criteria. In this context, an embedded shear layer develops along the outer region of the TBL leading to the formation of two-dimensional rollers typical of a Kelvin–Helmholtz instability which are captured by a spectral proper orthogonal decomposition (SPOD) analysis. The most energetic SPOD spatial modes of the tangential velocity show that streaks form along the airfoil suction side and, as the APG becomes stronger, they grow along the spanwise and wall-normal directions, having a spatial support along the entire boundary layer.

Improved self-adaptive turbulence eddy simulation for complex flows and stall prediction using high-order schemes

It is challenging to apply numerical simulations to accurately predict the stall behavior of aircraft equipped with high-lift devices. Simulations with Reynolds-Averaged NavierStokes (RANS) models suffer from lack of the reliability at high angles of attack with separated and reattached boundary layers, whereas wall-resolved Large Eddy Simulations (LES) of wall-bounded flows at high Reynolds numbers currently costs too much computational resources. A new unified hybrid turbulence modeling approach, denoted the Self-Adaptive Turbulence Eddy Simulation (SATES), is proposed and applied for complex turbulent flows combining with high-order numerical scheme of the Weighted Compact Nonlinear Scheme (WCNS) in the present study. It enables a seamless evolution from unsteady RANS to LES and finally approaches Direct Numerical Simulation (DNS) depending on the turbulent scales. In the framework of SATES, a new SATES-σ model with an adaptive model coefficient is developed by extending the underlying LES mode based on an enhanced sub-grid-scale model of the σ-model. The new SATES-σ is first examined in two benchmark cases of channel flow and flow past a square cylinder. Then, it is validated in supercritical flow past a circular cylinder to assess the performance of turbulent models. The results show significant improvements over the previous SATES and IDDES in the predictions of boundary layer flow. Finally, successful application is achieved in the accurate prediction of the stall of the MD-30P30N airfoil at a Reynolds number of 9×106 with wide angles of attack. The simulation results show good agreement with experimental results for surface pressure even for the challenging cases of 21 and 23 deg angles of attack. Again, the SATES-σ shows better results than the previous SATES and IDDES. The presented method has considerable potential for the challenging stall predictions.

Machine Learning Analysis of Thermal Performance Indicator of Heat Exchangers with Delta Wing Vortex Generators

In this work, the design features of delta wing vortex generators (DWVGs) on the thermo-hydraulic performance of heat exchangers are investigated using machine learning. Reynolds numbers, attack angle, length, wing-to-width ratio, and relative pitch ratio of DWVGs were used as descriptor variables, with Nusselt numbers, friction factors, and performance evaluation criterion (PEC) serving as target variables. Decision tree classification revealed the pathways leading to high or low values of the performance variables. Among many of those pathways, it was found that high Reynolds numbers (between 8160 and 9800) and high attack angles (greater than or equal to 47.5°) lead to high Nusselt numbers. On the other hand, an attack angle between 41° and 60°, a Reynolds number less than 8510, and a wing-to-width ratio greater than or equal to 0.4 causes a high friction factor. Finally, the PEC is likely to enhance when the Reynolds number is higher than or equal to 10,300 and the attack angle is between 47.5° and 60°. In addition to the decision tree analysis, SHapley Additive exPlanations (SHAP) analysis (a part of explainable machine learning) was also applied to reveal the importance of design features and their positive and negative effects on the target variables. For example, for a Nusselt number as the target variable, the Reynolds number was found to be the most influential variable, followed by the attack angle and the relative pitch ratio, all of which had a positive impact on the target. It was then concluded that machine learning methods could help provide strong insights into the configuration design features of heat exchangers in DWVGs to improve their efficiency and save energy.

Numerical study on aerodynamic and noise performance of bionic asymmetric airfoil with surface grooves

Based on the asymmetric NACA4412 baseline airfoil, a bionic airfoil with surface grooves is presented. For the bio-inspired airfoil, non-smooth grooves are placed on the trailing edge of NACA4412 airfoil. To reveal the effects of non-smooth structures of the trailing edge on the aerodynamic and noise performance of airfoil, large eddy simulation and Ffowcs Williams–Hawkings acoustic analogy are adopted to investigate the aerodynamic performance and acoustic characteristics of the baseline NACA4412 airfoil and bionic airfoil at the chord-based Reynolds number, Re = 1.2 ×105. The numerical results show that the aerodynamic performance of the bionic airfoil is better than that of the baseline airfoil when the angle of attack is 14°. For all the sound frequencies studied in this study, the overall sound pressure level of the bionic airfoil is reduced by 2.0 dB at angle of attack is 14°. At the same time, the mechanisms of flow control and noise reduction of non-smooth surface grooves at the trailing edge are also revealed. As a result, the presence of surface grooves near the trailing edge of the airfoil can effectively improve the aerodynamic performance and reduce the aerodynamic noise of the traditional asymmetric airfoil, especially at high angles of attack.

Experimental Propeller Performance Analysis of Distributed, Single and Isolated configurations

Distributed propulsion (DP) configurations are a promising concept for future aircraft systems. The main objective of the presented experiment is to investigate aerodynamic interactions of such configurations in detail and compare DP results to a single propeller configuration. The experimental setup at the Propulsion Test Facility, TU Braunschweig, features three co-rotating propellers. These are not attached directly to the wing, but are mounted on a separate carrier. This decoupling allows the forces and moments acting on wing and propeller to be considered separately. Additionally, different relative propeller positions are set up easily. In order to eliminate side wall effects, only the centre propeller and the centre wing element are subject of investigation for the distributed configuration. The periodically repeating outboard propellers reduce the wind tunnel interference while providing a true DP setup for the instrumented centre. Additionally to the DP setup, tests for conventional propeller wing configurations (only centre propeller installed) as well as isolated propeller and clean wing tests were performed. Thus, the DP effects on wing and propeller and the effect of the downstream wing on the propeller are clearly identified. The comparison between single propeller and distributed propulsion configurations shows that with distributed propulsion the drag increase is reduced from 326% to 216% compared to a clean wing. This effect is intensified by a greater thrust level and higher angles of attack. In order to identify the distributed propulsion effects, the forces acting on the propeller as well as the resulting efficiency of the propeller are compared between the distributed and single propeller configuration for two different relative propeller positions. The efficiency of the centre propeller is increased due to the outer propellers by approx. 2% to 3%.

Event-Based Guidance and Incremental Control with Application to Fixed-wing Unmanned Aerial Vehicle Perched Landing Maneuvers

Considering the nonlinearity and unknown dynamics of fixed-wing unmanned aerial vehicles in perched landing maneuvers, an event-based online guidance and incremental control scheme is proposed. The guidance trajectory for perched landing must be dynamically feasible therefore an event-based trapezoidal collocation point optimization method is proposed. Introduction of the triggering mechanism for the rational use of computing resources to improve PL accuracy. Furthermore, a filter-based incremental nonlinear dynamic inverse (F-INDI) control with state transformation is proposed to achieve robust trajectory tracking under high angle of attack (AOA). The F-INDI uses low-pass filters to obtain incremental dynamics of the system, which simplifies the design process. The state transformation strategy is to convert the flight-path angle, AOA and velocity into two composite dynamics, which avoids the sign reversal problem of control gain under high AOA. The stability analysis shows that the original states can be controlled only by controlling the composite state. Simulation results show that the proposed scheme achieves high perched landing accuracy and a reliable trajectory tracking control.

High propulsive performance by an oscillating foil in a stratified fluid

We investigate numerically the propulsion characteristics of an oscillating foil undergoing coupled heave and pitch motion in a linearly density-stratified flow. A parameter space defined by the internal Froude number ( $1 \le Fr \le 10$ ) and the maximum angle of attack ( $5^\circ \le {\alpha _0} \le 30^\circ$ ) is considered in our study. The results demonstrate a significant enhancement in both thrust production and propulsive efficiency due to the stratification influence. Notably, the highest efficiency exceeding $80\,\%$ is achieved under moderate stratification conditions, surpassing the performance observed in a homogeneous fluid. We attribute this optimum performance to the proper match between the stratification effect and foil kinematics, which gives rise to intense vortex interactions and sufficient wave–mean flow interactions in the near wake of the oscillating foil. Consequently, the energy is transferred towards wake structures to form a high-intensity momentum jet in close proximity to the foil's trailing edge, indicating efficient propulsion. Furthermore, we find that the stratifications within the moderate-to-strong transitional regime display a reduced dependence of propulsive efficiency on the maximum angle of attack, primarily due to the delaying and alleviating effects on dynamic-stall events. Such a mechanism enables the oscillating foil to maintain a satisfactory performance by sufficiently high angles of attack without the penalty of stall events. Based on our findings, we propose that animals or artificial vehicles utilising oscillatory propulsion can benefit from the presence of density stratification in the surrounding fluid.

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.

Effect of Vortex Generations on Aerodynamic Performance of Rough Wind Turbine Airfoil

Experimental investigations were carried out in wind tunnel facilities to assess the effect of vortex generators on the aerodynamic characteristics of a NACA4418 airfoil, particularly in scenarios where surface roughness is precipitated by environmental elements like dust, sand, and insects. These elements are known to adversely affect the efficiency of power generation. These studies meticulously analyzed the airfoil’s pressure distribution and its aerodynamic coefficients, taking into account the roughness positioned at varying locations along the chord. Special attention was given to the roughness located at 0.2c due to its pronounced impact on aerodynamic efficiency. The research highlighted that vortex generators play a crucial role in mitigating boundary layer separation, thereby enhancing the lift-to-drag ratio, especially at higher angles of attack. Notably, the deployment of vortex generators on airfoils affected by surface roughness resulted in a more pronounced improvement, elevating the maximum lift coefficient by 11.6% compared to their effect on smooth airfoils. Furthermore, the study revealed that an increase in the Reynolds number significantly augments the benefits provided by vortex generators, improving the lift coefficient and delaying the stall angle for airfoils with surface roughness, while also shifting the boundary layer separation point towards the airfoil’s trailing edge.

Investigations on bifurcation behavior of wind turbine airfoil response at a high angle of attack

Design load and vibration for parked conditions are gaining in importance for large-scale modern wind turbines with increasing flexibility, especially edgewise vibration when the blade is at a high angle of attack. In this work, flow-induced vibration of the wind turbine airfoil at 90 degrees of attack angle is studied with the fluid-structure interaction (FSI) simulation. The unsteady aerodynamic force due to flow separation and vortex shedding at the high angle of attack causes the chordwise vibration of the airfoil. When the vortex shedding frequency fv gets close to the chordwise natural frequency fn of the airfoil, vortex-induced vibration (VIV) of high amplitude occurs accompanied with the frequency lock-in phenomenon. In the post lock-in regime, it is found that period-3 and torus bifurcation occur successively and the vibration response becomes aperiodic. Dynamic mode decomposition(DMD) technique is used to investigate the mechanism of bifurcation from the perspective of energy balance, through analyzing the vorticity field in the wake and pressure distribution on the airfoil surface. For the certain incoming velocity in the post lock-in regime, since the frequency of the DMD mode f=2fv/3 is close to the natural frequencyfn, both the vibration of frequency 2fv/3 and fv get excited, leading to the onset of bifurcation. The Lissajou curves are obtained through reconstructing the transient pressure of each DMD mode, which indicates that energy transfer mainly exists in modes f=fv. In addition, the reconstructed Lissajou curves based on the leading DMD modes agree well with the original time-domain Lissajou curves.

Analysis of stall characteristics of the cuspidal waverider at subsonic conditions

The stall characteristics of the cuspidal waverider under subsonic conditions were studied at M = 0.4 and H = 0 km incoming flow. The stall characteristics were compared with those of delta wing models with the same leading-edge swept angle and the same aspect ratio. It can be found that the flow field changes of the cuspidal waverider are similar to those of Model 1 and Model 2. The critical angle of attack of the cuspidal waverider is more influenced by the aspect ratio, but less related to the leading-edge swept angle. The lift change of the cuspidal waverider is closely related to the vortex structure on the leeward side. As the angle of attack increases, the vortex structure becomes thicker and stronger, and moves toward the middle of the model under the action of the overflowing high pressure gas on the lower surface. Under the same angle of attack, the vortex structure under the effect of the adverse pressure gradient, the intensity along the flow direction decreases while the vortex continues to move away from the wall. Besides, the possible splitting of large vortex into smaller ones will be another result of the effect of the inverse pressure gradient. The smaller the overall aspect ratio of the model, the greater actual swept angle and the greater effective swept angel of the vortex structure, which allows the vortex to exist at higher angles of attack. When the vortex structure is broken to form a larger recirculation area in the leeward area, the negative pressure at the leeward surface decreases, the lift coefficient decreases, and a stall occurs.

Aircraft Upset Recovery Strategy and Pilot Assistance System Based on Reinforcement Learning

The upset state is an unexpected flight state, which is characterized by an unintentional deviation from normal operating parameters. It is difficult for the pilot to recover the aircraft from the upset state accurately and quickly. In this paper, an upset recovery strategy and pilot assistance system (PAS) based on reinforcement learning is proposed. The man–machine closed-loop system was established and the upset state, such as a high angle of attack and large attitude angle, was induced. The upset recovery problem was transformed into a sequential decision problem, and the Markov decision model of upset recovery was established by taking the deflection change of the control surface as the action. The proximal policy optimization (PPO) algorithm was selected for the strategy training. The adaptive pilot model and the reinforcement learning method proposed in this paper were used to make the aircraft recover from the upset state. Based on the correspondence between the flight state, the recovery method, and the recovery result, the aircraft upset recovery safety envelopes were formed, and the four-level upset recovery PAS with alarm warning, coordinated control, and autonomous recovery modes was constructed. The results of the digital virtual flight simulation and ground flight test show that compared with a traditional single pilot, the aircraft upset recovery strategy, the upset recovery safety envelopes, and the PAS established in this study could reduce the handling burden of the pilot and improve the success rate and effect of upset recovery. This research has certain theoretical reference values for flight safety and pilot training.

Flow Separation Control of an Airfoil by Using External Fluid Flow at a High Angle of Attack

Flow Separation Control of an Airfoil by Using External Fluid Flow at a High Angle of Attack

Longitudinal Attitude Control and Stability Analysis for a Low Aspect Ratio Flying Wing UAV at High Angle of Attack

Longitudinal Attitude Control and Stability Analysis for a Low Aspect Ratio Flying Wing UAV at High Angle of Attack

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.