High Angles Of Attack Research Articles

Research on Minimum Unstick Speed Flight Test Method for Civil Aircraft

The minimum unstick speed (VMU) is defined as the calibrated airspeed at which an aircraft can safely lift off and continue its takeoff without any difficulties. The VMU flight test, characterized by low-speed and high angle of attack under the influence of ground effect, poses significant challenges and high risks, making it one of the most hazardous flight test subjects in civil aircraft flight test. The research on VMU flight test methods for civil aircraft is of utmost importance in strengthening the foundation of flight test technology, improving flight test efficiency, and ensuring flight test safety.

Flow Separation Control and Aeroacoustic Effects of a Leading-Edge Slat over a Wind Turbine Blade

To enable wind energy to surpass fossil fuels, the power-to-cost ratio of wind turbines must be competitive. Increasing installation capacities and wind turbine sizes indicates a strong trend toward clean energy. However, larger rotor diameters, reaching up to 170 m, introduce stability and aeroelasticity concerns and aerodynamic phenomena that cause noise disturbances. These issues hinder performance enhancement and social acceptance of wind turbines. A critical aerodynamic challenge is flow separation on the blade’s suction side, leading to a loss of lift and increased drag, ultimately stalling the blade and reducing turbine performance. Various active and passive flow control techniques have been studied to address these issues, with passive techniques offering the advantage of no external energy requirement. High-lift devices, such as leading-edge slats, are promising in improving aerodynamic performance by controlling flow separation. This study explores the geometric parameters of slats and their effects on wind turbine blades’ aerodynamic and acoustic performance. Using an adequate turbulence model at Re = 106 for angles of attack from 14° to 24°, 77 slat configurations were evaluated. Symmetric slats showed superior performance at high angles of attack, while slat chord length was inversely proportional to aerodynamic improvement. A hybrid method was employed to predict noise, revealing slat-induced modifications in eddy topology and increased low- and high-frequency noise. This study’s main contribution is correlating slat-induced aerodynamic improvements with their acoustic effects. The directivity reveals a 10–15 dB reduction induced by the slat at 1 kHz, while the slat induces higher noise at higher frequencies.

Enhancement of flying wing aerodynamics in crossflow at high angle of attack using dual synthetic jets

To address the asymmetric flow field of the flying wing under cross-flow conditions at high angles of attack, dual synthetic jet actuators (DSJAs) are positioned at the leading edge of the windward side. DSJ control effectively affects the flow separation structure on the windward side, thereby enhancing leading-edge suction. This leads to both lift enhancement and drag reduction. Furthermore, it strengthens the stabilizing roll moment and improves lateral static stability. Additionally, the effective suppression of the separation zone significantly reduces the aerodynamic load fluctuations of the flying wing after control. In terms of excitation frequency, low-frequency excitation more effectively generates lift, transfers energy downstream, promotes momentum transfer, and results in an 8.2% increase in lift. Moreover, the DSJ excitation exhibits fast response characteristics, with the entire flow establishment process taking only 190 ms. Therefore, through DSJ control, the asymmetric flow under lateral conditions can be effectively corrected, expanding the flight envelope and enhancing the maneuverability and reliability of the aircraft.

Unsteady Propeller Wake Interference on Wing in Tractor Configuration at Low-Reynolds-Number Condition

This study investigates the unsteady interaction between the propeller slipstream and NACA0012 airfoil, which causes laminar separation at a moderate angle of attack and nonlinear aerodynamic characteristics at low Reynolds numbers. The Reynolds number based on chord length was set to c, and the propeller advance ratio was set to J=0.8. Wind-tunnel experiments were conducted, including aerodynamic force measurement and time-resolved particle image velocimetry visualization. Regarding time-averaged characteristics, the propeller installed condition exhibits a linear change in lift; reduced drag; and, consequently, an improved lift-to-drag ratio. Consistent with the aerodynamic characteristics, suppressed separation and reduction in wake deficit are confirmed in the time-averaged flowfields. Additionally, vortex formation and shedding synchronized with propeller rotation are observed in unsteady flowfields at the middle and high angles of attack. This periodic and large-scale vortex structure suggests the possibility of further improving the performance of the propeller–wing system in low Reynolds numbers.

CFD Investigation of Dual Synthetic Jets on an Optimized Aerofoil's Trailing Edge

In fluid dynamics, a flow control device is used to control, manage, or modify the behavior of a fluid flow. Jet actuators work by releasing high-velocity jets of fluid, usually air or gas, into the surrounding environment to control or manipulate the flow of fluids. In this study, the flow control device, which was a dual synthetic jet actuator (DSJA), acted as a lift enhancement device over an optimized NACA 0012 aerofoil with a rounded trailing edge (TE) (Coanda surface approximately 9% of the trailing edge was modified) to enhance the lift at various angles of attack (AOAs). Fluctuating pressure inlets were introduced in two slots. When the dual synthetic jets were in control, the out-of-phase jets from the upper and lower trailing edge jets helped to boost the lift coefficient. The suction stroke from the lower half of the jet made the Coanda effect stronger in the upper half. The upper trailing edge jet deflected downwards merged with the lower one and helped to deflect the flow field closer to the bottom half. An unsteady CFD analysis was performed on optimized airfoils with and without a DSJ, with a driving frequency of 40.6 and a reduced frequency of 0.025 at a Reynolds number of 25000. The results obtained at different angles indicated that the L/D ratio was improved by 13.5% at higher angles of attack in the presence of the DSJA.

Design of Input Signal for System Identification of a Generic Fighter Configuration

This article investigates the design of time-accurate input signals in the angle-of-attack and pitch rate space to identify the aerodynamic characteristics of a generic triple-delta wing configuration at subsonic speeds. Regression models were created from the time history of signal simulations in DoD HPCMP CREATETM-AV/Kestrel software. The input signals included chirp, Schroeder, pseudorandom binary sequence (PRBS), random, and sinusoidal signals. Although similar in structure, the coefficients of these regression models were estimated based on the specific input signals. The signals covered a wide range of angle-of-attack and pitch rate space, resulting in varying regression coefficients for each signal. After creating and validating the models, they were used to predict static aerodynamic data at a wide range of angles of attack but with zero pitch rate. Next, slope coefficients and dynamic derivatives in the pitch direction were estimated from each signal. These predictions were compared with each other as well as with the ONERA wind tunnel data and some CFD calculations from the DLR TAU code provided by the NATO Science and Technology Organization research task group AVT-351. Subsequently, the models were used to predict different pitch oscillations at various mean angles of attack with given amplitudes and frequencies. Again, the model predictions were compared with wind tunnel data. Final predictions involved responses to new signals from different models. A feed-forward neural network was then used to model pressure coefficients on the upper surface of the vehicle at different spanwise sections for each signal and the validated models were used to predict pressure data at different angles of attack. Overall, the models predict similar integrated forces and moments, with the main discrepancies appearing at higher angles of attack. All models failed to predict the stall behavior observed in the measurements and CFD data. Regarding the pressure data, the PRBS signal provided the best accuracy among all the models.

Numerical simulation of cavitation flow around a wing with new tubercles design

The study investigates the benefits of adding non-uniform leading-edge tubercles to hydrofoils, drawing inspiration from the complex shaping of humpback whale flippers. The focus is on managing cavitation, a phenomenon that can affect hydrofoil performance. While previous research has mainly looked at uniform sinusoidal tubercles and their positive impact on flow dynamics and cavitation control, this study introduces a new perspective by examining the effects of non-uniform tubercles on hydrofoil performance. Advanced computational fluid dynamics (CFD) simulations using ANSYS Fluent 2024 were used to assess how these non-uniform tubercles affect the lift, drag, and cavitation characteristics of the NACA 634-021 hydrofoil. The simulations incorporated the Schnerr-Sauer cavitation model and the SST k-ω turbulence model to accurately capture the flow dynamics. The results show that non-uniform tubercles improve cavitation control by disrupting the flow in a way that delays the onset and reduces the severity of cavitation. The modified hydrofoils (non-uniform tubercles) display improved hydrodynamic performance compared to baseline designs, with significant reductions in drag and increased lift at higher angles of attack. This study provides valuable insights into the potential of mimicking natural designs to enhance flow stability and address cavitation issues, offering a significant contribution to advanced hydrofoil design in scenarios where controlling cavitation is essential. The broader implications of this research underscore the potential of mimicking natural designs to transform hydrofoil engineering and enhance flow stability in various applications.

Analysis of aerodynamic characteristics of drone wing based on CFD

In order to improve the aerodynamic characteristics of the drone wings, CFD method was used to simulate and calculate the lift coefficient, drag coefficient, and lift-drag ratio under different relative inflow velocities, as well as the velocity and pressure fields under different attack angles. Modal calculations were conducted on the wing to obtain the first four modal shapes, providing a basis for analyzing flutter characteristics. An iterative calculation method of incompressible potential flow-boundary layer based on surface element method was combined with the software XFOIL to optimize the airfoil at low wind speeds. The results indicate that the airfoil is susceptible to stall at high angles of attack, with the pressure of the separation flow being nearly equivalent to that at the separation point. Subsequent to separation, there is an increase in differential pressure resistance, resulting in a marked rise in the drag coefficient. At the optimized angle of attack, the lift-drag ratio of the optimized wing increases by 12.58 %, while there is a decrease of 0.084 % in lift coefficient and an increase of 11.21 % in drag coefficient.

Development on Unsteady Aerodynamic Modeling Technology at High Angles of Attack

Development on Unsteady Aerodynamic Modeling Technology at High Angles of Attack

Optimal Selection of Active Jet Parameters for a Ducted Tail Wing Aimed at Improving Aerodynamic Performance

The foldable tail of the box-type launch vehicle poses a risk of mechanical jamming during the launch process, which is not conducive to the smooth completion of the flight mission. The integrated nonfolding ducted tail proposed in this article can solve the problem of storing the tail in the launch box. Moreover, traditional mechanical control surfaces have been eliminated, and active jet control has been adopted to control the pitch direction of the flight attitude, which can improve the structural reliability of the tail wing. By studying the effects of parameters such as momentum coefficient, jet hole position, jet hole height, and jet angle on improving the aerodynamic performance of ducted tail wing, relatively good jet parameters are selected. Research has found that compared with jet hole height and jet angle, momentum coefficient and jet hole position are more effective in improving the aerodynamic performance of ducted tail wings. Under a trailing edge jet, a relatively good jet condition occurs when the jet hole height is equal to0.25% of the aerodynamic chord length, and the jet angle is equal to 0°. At this time, with the increase of the jet momentum coefficient, the effect of increasing the lift of the ducted tail wing is the best. Finally, a comparative analysis is conducted on the lift and drag characteristics between the ducted tail wing and traditional tail wing, and it is found that the ducted tail wing can generate lift at a 0° attack angle and will not stall in the high attack angle range of 12°~22°, with broad application prospects.

Research on assisted recovery algorithm of stall state based on signal guidance

Abstract Given the lack of quantitative guidance in the current stall recovery process, which leads to the pilot’s misoperation or untimely recovery, referring to the energy control theory of complex state recovery, a stall state-assisted recovery algorithm based on pitch angle, engine thrust, and roll angle signal guidance was determined. The structure pilot model and the kinematics simulation model of aircraft at a high angle of attack were established. The effectiveness of the algorithm has been verified by simulation analysis. The man-machine closed-loop simulation analysis results show that the algorithm can guide the pilot in effectively recovering from stall conditions under different stall scenarios and engine failure conditions.

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.

Omnidirectional control of the wake of a circular cylinder with spinning rods subject to a turbulent flow

We numerically investigate the attribute of omnidirectionality of the flow-control system comprised of a large circular cylinder equipped with eight spinning rods of smaller diameter, subject to an incoming flow that adopts different angles of attack. Detached-eddy simulations are employed to compute hydrodynamic loads and to provide flow topology at a Reynolds number of 1000. Two cases are assessed regarding the rods angular velocities. In case 0, all rods spun with the same angular velocity. In case 1, velocities were inspired by potential-flow theory. The two systems have the same input kinetic energy in common. To assess the system response, the velocities were increased proportionally. Both cases succeeded in reducing the mean drag. However, while case 1 proved to become ever “more omnidirectional” with increasing angular velocities, case 0 demonstrated to be prone to the angle of attack as it was unable to suppress vortex shedding for sufficiently large slopes of the incoming flow, and in such circumstances, unable to reduce hydrodynamic forces. We verify that the lift is mitigated in case 1, in contrast to case 0. Even for a vortex-free downstream flow resulting from configurations of high velocities and high angle of attack, the latter produces asymmetric recirculation regions downstream of the system that drive a pressure imbalance. The different outcomes of the two systems are also explored from the viewpoint of power consumption, and it is revealed that the omnidirectionality of case 1 is intrinsically related to the emphasis imposed on rotation rates of a subset of the eight rods.

Design perspectives of a system identification approach of the NATO AVT-316 multi-swept wing aerodynamics

Air supremacy requires enhanced maneuvering capabilities at high angles of attack and even beyond the stall angle. High angle-of-attack (high-α) maneuverability can be achieved using swept wings and tailored leading-edge shapes to replace local and disorganized flow separation with controlled separation-induced leading-edge vortices (LEV). Strong leading-edge vortices delay the stall to higher angles and generate a non-linear lift increment up to the angle of attack where the jet-type flow structure of LEV changes to a reversed-flow bubble (vortex breakdown phenomenon). A multi-swept wing configuration has the potential to delay the vortex breakdown to even higher angles as the vortices formed over the front wing can energize vortices formed over the main wing and lead to vortex merging. This article is focused on understanding the unsteady interactions between multiple leading edge vortices formed over multi-swept wing configurations in the subsonic speed regime. Specifically, this article investigates the aerodynamic characteristics of wings being evaluated under the initiative of the NATO STO AVT-316 Task Group. Highly refined meshes, and the use of hybrid turbulence models such as Detached Eddy Simulations (DES) and Delayed DES are required to accurately resolve the vortical flows over these wings. Simulating all flow conditions of interest is a computationally expensive approach. A system identification method was therefore proposed to rapidly and accurately generate aerodynamic models of these wings. The method uses a novel piece-wise chirp (constant amplitude and increasing frequency) motion as an input signal (training maneuver). The motion computational cost is equivalent to the cost of six static CFD simulations, however it can predict aerodynamic responses of the wings over a wide range of angles of attack. The method has been tested for double and triple delta wings. Some design considerations are provided based on predicted flow features and aerodynamic data. Prediction results with different turbulence models, sting geometries, grid resolutions and an adaptive mesh refinement approach are provided.

Aerodynamic Analysis and Simulation for a Z-Shaped Folding Wing UAV

The potential of folding wing unmanned aerial vehicles (UAVs) in enhancing the adaptability of complex missions is substantial. Analyzing the aerodynamic characteristics of these UAVs is vital for optimizing the design of their folding wing configuration and airfoil shape. In this study, a model of a Z-shaped folding wing UAV was established with the objective of enhancing UAV maneuverability. An integrated vortex lattice method (VLM) was employed to analyze the aerodynamic characteristics of the Z-shaped folding wing and numerical simulations were utilized to validate the obtained results, which prove the effectiveness of the integrated VLM in studying aerodynamic characteristics at a high angle of attack of the Z-shaped folding wing UAV. This is a new way to reduce computation time while maintaining accuracy to calculate such complex folding structures in high fidelity. Furthermore, the Z-shaped folding wing configuration demonstrates enhanced maneuverability at high angles of attack, and a Simulink flight simulation model based on the aerodynamic coefficients of the Z-shaped folding wing verifies this property. This analysis can properly predict the post-stall characteristics for the Z-shaped folding wing UAV to ensure the safety of the vehicle during flight.

Optimizing Aerofoil Design: A Comprehensive Analysis of Aerodynamic Efficiency through CFD Simulations and Wind Tunnel Experiments

This study explores the aerodynamic properties of different aerofoil shapes and their performance under varying flow conditions to identify the most efficient design based on lift-to-drag ratio, stall behaviour, and overall aerodynamic efficiency. Using Computational Fluid Dynamics (CFD) simulations, several aerofoil profiles were analysed at different angles of attack and flow speeds. These simulations were validated through wind tunnel experiments, offering a comprehensive understanding of aerofoil performance in real-world scenarios. The combination of CFD analysis and wind tunnel testing enabled a thorough assessment of each aerofoil shape, leading to the discovery of a specific aerofoil with a high lift-to-drag ratio and stable performance at high angles of attack. These results have significant implications for the design of wings and blades in aerospace and aeronautical applications, improving fuel efficiency and performance in both aviation and wind energy sectors. Additionally, dynamic roughness shows potential in reducing separation bubbles, but further investigation is needed to assess its effectiveness at higher angles of attack and elevated Reynolds numbers. Understanding the scalability and practical application of dynamic roughness in real-world scenarios is essential. Current research on surface modifications like dimples and riblets lacks optimized configurations for varying conditions. More research is needed to understand the interaction between surface geometries and the boundary layer, particularly at higher angles of attack and Reynolds numbers. Combining experimental and numerical methods can provide a comprehensive understanding of flow control techniques. The limited research on applying flow control strategies to wind turbine blades indicates a significant opportunity to improve wind energy efficiency. Future studies should focus on optimizing multiple techniques and addressing practical challenges, such as durability, cost-effectiveness, and integration into existing systems. Investigating the cost-effectiveness and durability of these modifications for long-term use will be vital for their successful adoption in the industry. Expanding research to include the effects of environmental factors like temperature and humidity will offer a more complete understanding of flow control in various operating conditions. By addressing these gaps, advancements in aerodynamic performance can be achieved, benefiting the aerospace and wind energy sectors.

Flight Stability of Spinning Projectiles at High Angles of Attack

The flight behavior of the spin-stabilized 105 mm M1 artillery projectile, when fired with high quadrant elevation, is studied using six-degrees-of-freedom trajectory simulations and theoretical methods of stability analysis. Trajectory simulation results confirm the known effect of overstabilization at the apogee, leading to high angles of attack (AoA), drift to the left, and base-forward descent. For quadrant elevations between nose-forward and base-forward descent, a previously unknown descent mode was discovered, with projectiles performing a low-frequency coning motion at high angles of attack of 73 and 104 deg, respectively. An exhaustive theoretical description of such high-AoA limit cycles is given that overcomes the limitation to small angles of attack prevalent in traditional concepts of projectile stability. Based on a formulation of the angular motion in spherical coordinates, conditions for the existence of limit cycles are identified and stability criteria are derived using Lyapunov’s first method of linearization around equilibrium and eigenvalue analysis of the resulting Jacobian. The resulting criteria indicate that limit cycle flight is primarily governed by the Magnus moment coefficient. The method is applied to the 105 mm M1 projectile, and the coning motions observed in the trajectory simulations are accurately reproduced by the theory.

HEMLAB Algorithm Applied to the High-Lift Japan Aerospace Exploration Agency Standard Model

The high-lift Japan Aerospace Exploration Agency (JAXA) standard model (HL-JSM) has been numerically analyzed in order to further validate the HEMLAB code for realistic aircraft configurations. The numerical algorithm is based on highly efficient edge-based data structure for a vertex-based finite volume algorithm on hybrid meshes. The data pattern is arranged to meet the requirements of a vertex-based finite volume algorithm by considering data access patterns and cache efficiency. A fully implicit version of the numerical algorithm has also been implemented based on the Portable, Extensible Toolkit for Scientific Computation (PETSc) library in order to improve the robustness of the algorithm. The resulting algebraic equations, including the one-equation Spalart–Allmaras, are solved in a monolithic manner using the restricted additive Schwarz preconditioner combined with the flexible generalized minimal residual method [FGMRES(m)] Krylov subspace algorithm. The numerical method is also combined with the metric-based anisotropic mesh refinement library Python Adaptive Mesh Geometry Suite–INRIA (pyAMG) from National Institute for Research in Digital Science and Technology in order to improve the numerical accuracy. The numerical algorithm is initially applied to the two-dimensional L1T2 (National High Lift Programme) high-lift system, and then the calculations around the HL-JSM are carried out at relatively high angles of attack. The numerical results with the anisotropic mesh refinement library pyAMG indicate significant improvements in numerical accuracy.

Study on the influence of high-temperature thermochemical nonequilibrium effects on the flow field characteristics around shaped wing

With advances in research on hypersonic vehicles, the precise simulation of the effects of thermochemical non-equilibrium has become increasingly important in their design. In light of this, this study explores the influence of high-temperature thermochemical non-equilibrium on the characteristics of the flow field around the hypersonic wings of an aircraft. We initially conducted a numerical simulation by using the model of flow through a cylinder to validate the accuracy and reliability of an 11-species gas model in representing high-enthalpy flow fields. Subsequently, a systematic analysis was conducted on the impact of thermochemical nonequilibrium effects on the temperature, pressure, and enthalpy distribution in the flow field around a symmetric diamond wing under different Mach numbers and angles of attack. The research results indicated that the deeper reason behind the differing thermochemical nonequilibrium effects in the flow field at various Mach numbers lies in the distinct distribution of enthalpy of the air components at different locations, which provided a new perspective for understanding flow field variations from the standpoint of enthalpy. It is disclosed that the thermochemical non-equilibrium significantly altered the characteristics of the flow field, particularly at high Mach numbers and angles of attack, with a significant impact on the aerodynamic parameters of both the windward and the leeward sides of the wing. Through explaining the mechanisms of thermochemical non-equilibrium under flow fields with different structures, this study provides a theoretical foundations for and a fresh perspective on the design of hypersonic vehicles.

Flight Dynamic Uncertainty Quantification Modeling Using Physics-Informed Neural Networks

When attempting to develop aerodynamic models for extreme flight conditions, including high angle of attack, high rotational rates, and tumbling motion, many classical methods have challenges in accurately modeling the highly nonlinear aerodynamic effects present. Physics-informed neural networks (PINNs) have previously been shown to be a potential technique to model these nonlinear aerodynamic effects when framed as a system identification problem. PINNs are well suited to this problem as they benefit from the universal approximation abilities of neural networks while directly incorporating known physical constraints into the training process. One of the main challenges in machine learning algorithms, including PINNs, is quantifying the confidence in a deterministic model prediction. This work expands on the previous development of PINNs as an aerodynamic and system identification tool by incorporating uncertainty quantification through three ensemble methods to provide calibrated confidence intervals on both aerodynamic coefficients and propagated trajectories. This work demonstrates and evaluates these methods on a simulated F16 case study where up to 100 PINN models are trained on varying training datasets. These models provide aerodynamic coefficients directly and are used to propagate trajectories within a six-degree-of-freedom simulation environment.