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With high aerodynamic efficiency and low observable design for flying wing configuration, the control ability decreases and the flying risk increases, especially during the takeoff and landing processes for unmanned aerial vehicle (UAV). To solve the nonlinear characteristic problem of the flying wing configuration at a high angle of attack (AoA) for a low aspect ratio flying wing UAV, a controllable region and stability analysis method is proposed in the paper. Firstly, the aerodynamic characteristics and static stability of a flying wing UAV are analyzed. Secondly, according to the characteristics of aerodynamic data, an efficient longitudinal trim state calculation method under the low‐speed range is proposed. The trim state contour map can be used effectively with the proposed trim state calculation method, which can further analyze the constraint relationship between different trim states and the controllable region of the UAV. Thirdly, the stability of different trim states is analyzed through the phase portrait between AoA and airspeed. With the analysis result, longitudinal attitude and AoA decoupling control laws are designed based on the incremental nonlinear dynamic inversion (INDI) method. The effectiveness and adaptability of the designed control methods are verified through simulation at high AoA. [ABSTRACT FROM AUTHOR]

Strong coupling of aerodynamics and dynamic motion will occur on advanced highly maneuverable aircraft at a high angle of attack, which makes the complex flow field heavily nonlinear. To improve the aerodynamic performance at a high angle of attack and broaden the flight boundary of the aircraft, a new combination configuration of a prepositive elliptical wing–main wing is proposed. The high-energy wake vortex generated by the stationary or vibrating prepositive wing is mixed with the airflow field of the main wing at the high angle of attack, and the air flow is accelerated through the space between the elliptical wing and its main wing, thereby enhancing the ability of the boundary layer to resist separation. A two-dimensional combined configuration of a stationary or vibrating prepositive elliptical NACA0012 airfoil was designed, and the flow control simulation was studied with a high-precision computational fluid dynamics (CFD) method. The results show that the stationary elliptical wing generates lift by itself, the generated wake suppresses the flow separation on the upper wing surface of the main wing, and the lift-to-drag ratio can be increased by as much as 319%. Moreover, when the prepositive wing vibrates at a small amplitude, the generated wake energy is stronger and the control effect on the separation of the main wing is much more remarkable. In the best case, the lift-to-drag ratio of the main wing can be increased by 485%. Compared with conventional flow control techniques, the new combined flow control method has the characteristics of a simple form and a high degree of control without forming waste resistance, and significantly can improve the flow separation characteristics at high angles of attack and enhance the maneuvering ability of aircraft. [ABSTRACT FROM AUTHOR]

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 R e = 4 × 1 0 5 , 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. • Mean velocity profiles with strong APGs depict three inflection points, the outer one being unstable under inviscid stability criteria. • The production of TKE is affected by additional terms besides the shear component, especially in the outer layer. • The turbulence anisotropy state changes from a TBL to a shear layer under the APG influence. • The most energetic SPOD mode of tangential velocity shows streaks which grow along the spanwise and wall-normal directions due to strong APGs. • In the outer region of the TBLs, Kelvin–Helmholtz structures are depicted by the most energetic SPOD mode of wall-normal velocity. [ABSTRACT FROM AUTHOR]

Summary: This article attempts to study the high angle of attack maneuver from the perspective of switched system control. In view of the complex aerodynamic characteristics, an improved longitudinal attitude motion model is presented, which is a switched stochastic nonstrict feedback nonlinear system with distributed delays. The significant design difficulty is the completely unknown diffusion and drift terms and distributed delays with all state variables. Based on a technical lemma and neural networks, an improved smooth state feedback control law for nonstrict feedback systems is proposed without any growth assumptions. To eliminate the influence of distributed delays, an improved Lyapunov–Krasovskii function is constructed, which skillfully removes the constraint of the upper bound of the delay change rate. Then, by combining the average dwell‐time scheme and stochastic backstepping technique, an adaptive neural network tracking control law is designed, which extends a newly proposed switched system stability condition to the stochastic switched system. Theoretical analysis and flight control simulation experiments are provided to illustrate the effectiveness of the proposed control method. [ABSTRACT FROM AUTHOR]

The dynamic stall at high angle of attack is an important aerodynamic problem faced by aircraft, and it has always been a hotspot of aerodynamic research. The traditional reduced order model (ROM) methods needs to establish an accurate model, and has a high demand for experience. In this paper, a novel nonlinear aerodynamic identification method based on recurrent neural networks (RNNs) is proposed. The computational fluid dynamics (CFD) method is used to calculate the unsteady aerodynamic parameters of the NACA0012 airfoil. A group of sinusoidal chirp signals with variable amplitude and frequency are adopted as the excitation signals, and the obtained data are used to train the recurrent neural networks, and the ROM of the nonlinear aerodynamic model of high angle of attack dynamic stall is obtained. Finally, the aerodynamic parameters of a group of composite sinusoidal motion signals different from the training signals are predicted by the trained neural networks model and compared with the CFD results. [ABSTRACT FROM AUTHOR]

This paper aims to investigate the side force on a cone-cylinder geometry at different angles of attack (α) by adopting experiments and computations. The cone-cylinder configuration had a length to diameter ratio (L/D) of 10, and a base diameter (D) of 25 mm. The nose shape had a fineness ratio of approximately 3. Results indicated that the side force increases with the increasing of the angle of attack. A circular ring was used to reduce the side force at different angles of attack. Using a smaller height ring (2% of local diameter) in the initial portion of the body did not reduce the side force significantly at lower angles of attack. However, a ring with larger height (5% of local diameter) placed at 2.5 times the diameter of the body from the tip reduced the side force at almost all the angles of attack. [ABSTRACT FROM AUTHOR]

Abstract The pattern of asymmetric vortices over a blunt-nose slender body is manageable and highly dependent on the location of artificial microperturbation (AMP) on the nose at a high angle of attack (AoA). A particle is introduced and attached to the nose surface to examine the mechanism of asymmetric flow induced by AMP at high AoA (α = 50°). Wind tunnel experiments and numerical simulations were conducted to investigate the mechanism of asymmetric flow generated over the blunt-nose slender body. Perturbation flow (PF), which is separated from the AMP surface, divides the shear layer separating from the nose of the slender body into two parts (long and short). The shear layer separating is shown as separation line. The long part on the opposite side with AMP rolls up a vortex earlier than the short one on the same side with AMP. Moreover, the vortex from the long part exhibits higher vorticity than that from the short one. The differences in the two parts result in asymmetric vortices over the blunt-nose body at high AoA. The PF develops into the opposite side with AMP, thereby increasing the asymmetry of flow. The associated flow separation and evolution of streamlines are discussed in detail. [ABSTRACT FROM AUTHOR]

The aerodynamic characteristics of a NACA0012 wing geometry at low Reynold's numbers and angle of attack ranging from 0° to 90° are investigated using numerical simulations and the results are validated by wind tunnel experiments. Further experiments are conducted at low Reynold's numbers of 1 × 105, 2 × 105 and 3 × 105. Findings of the study show a similar trend for the lift and drag coefficients at all the investigated Reynold's numbers. The lift coefficient is linearly increased with angle of attack until it reaches its maximum value at 32° which is the stall angle. It is observed that further increment in angle of attack results in decrement of lift coefficient until it reaches its minimum value at 90° angle of attack. The drag force acting on the airfoil increases as the angle of attack is increased and increment in the drag force results in change of laminar flow to turbulent flow. As the turbulence gets higher the flow starts to separate from the airfoil surface due to eddies generated by turbulence. Hence, the lift force generated by the wing is reduced and drag force is increased simultaneously, which results in poor performance of the wing. [ABSTRACT FROM AUTHOR]

The nominal aerodynamic parameters of aircraft are often approximate and aircraft may experience high value of angle of attack/sideslip perturbations during their manoeuvres. Preventing instability and air crash requires a robust controller capable of containing the dynamic uncertainty and the perturbations. In this respect, the problems of roll control in such a situation are studied and a better choice of sliding surface is proposed. Sliding mode control manages the uncertainty and adaptive fuzzy is employed to shape the transient response. As a result, a setup is formed which outperforms the basic controller both in terms of transient speed, response robustness and control effort. The strength of the method is more appreciated in case of high angle of attack/sideslip perturbed manoeuvres. This is proved theoretically and illustrated by simulations. [ABSTRACT FROM AUTHOR]

In this paper, an extended unsteady aerodynamic model for aircraft oscillations at a high angle of attack is developed and modified based on the traditional state space model. With an analysis of the nonlinear flow field in the dynamic process, three influencing factors, the pitching angular velocity, the reduced frequency and the amplitude, are determined and added to the original state space equation to provide a more comprehensive description of the dynamic flow field. The particle swarm optimization (PSO) algorithm is used to identify unknown parameters in the extended model based on Computational Fluid Dynamics (CFD) results or wind tunnel experimental data. The identifiability of the parameters and the performance of the extended model are validated with a 70° sharp leading-edge delta wing and a scaling geometry of an F-18 aircraft. The results show that the extended model can achieve more accurate aerodynamic loads than the original model. Moreover, the aerodynamic forces and moments of various configurations at different reduced frequencies and amplitudes can be predicted with the model. The extended state space model is demonstrated to be a more efficient and widely applicable approach for unsteady aerodynamics modeling. [ABSTRACT FROM AUTHOR]

Asymmetric vortices over its blunt-nose slender body are generated when a missile flies at high angles of attack (AoA, up to 40–50°) to attain high manoeuvrability. The pattern of the asymmetric vortices changes randomly, due to uncertainties in manufacturing and surrounding conditions, thus resulting in unexpected side-forces (in both magnitude and direction). In order to control the behaviour of the asymmetric vortices, a micro-perturbation is introduced and attached on the nose of the slender body at a high AoA α = 50°. The micro-perturbation is a hemispherical protrusion with a radius of r = 0 . 0015 − 0 . 015 D , where D is the diameter of the blunt-nose slender body. Experimental tests and numerical simulations are conducted to investigate disturbed flows over the blunt-nose slender body at a Reynolds number R e D = 1 . 54 × 1 0 5 , based on oncoming free-stream velocity and D . The side-forces, pressure distributions and vortical structures are measured using force balance, pressure scanning and particle-image velocimetry (PIV), respectively. Three-dimensional numerical simulations are conducted, with the shear-stress transport (SST) k − ω turbulence model employed. It is found that the randomness features of the asymmetric vortices pattern were inhibited in the presence of the tiny perturbation. The pattern of asymmetric vortices and the corresponding side-forces are manageable and highly dependent on the location and the size of the perturbation. The associated pressure distributions and flow separation are discussed in detail. [ABSTRACT FROM AUTHOR]

At high angles of attack, the conventional rudders of fighters or tactical missiles are affected by the wake flow generated by the fuselage, and its yaw control power may dropdramatically together with leeward asymmetry vortices inducing asymmetry pressure distributions, provoking a side force with the same order of magnitude as that of normal force, and accompanying a strong yaw moment. Forebody vortex control methods can provide a desired yaw control moment for aircrafts at high angles of attack, and they have wide application prospect in the field of maneuver flight at high angles of attack. In this paper, the new methods developed in the past ten years were reviewed. The passive methods include boundary layer transition strips, bumps, dimples, strakes, nose bluntness, fluttering flag, and vortex generators etc. The active methods include plasma actuators, micro blowing, nose blowing, synthetic jet, fast-swing mini-tip strake and microballoon array actuators etc. The control effect, mechanism and applicable scope were emphatically introduced for various methods. The maneuverability of slender body at high angles of attack can be improved by the linear control methods including vortex generators, synthetic jet, fast-swing mini-tip strake, plasma actuators, micro blowing, which are practical in engineering application. Finally, the application prospect and development trend were discussed for the forebody vortex control method at high angles of attack. [ABSTRACT FROM AUTHOR]

This paper studies low-aspect-ratio (...) rectangular wings at high incidence and in side-slip. The main objective is to incorporate the effects of high angle of attack and side-slip into a simplified vortex model for the forces and moments. Experiments are also performed and are used to validate assumptions made in the model. The model asymptotes to the potential flow result of classical aerodynamics for an infinite aspect ratio. The ... → 0 limit of a rectangular wing is considered with slender body theory, where the side-edge vortices merge into a vortex doublet. Hence, the velocity fields transition from being dominated by a spanwise vorticity monopole (... ≫ 1) to a streamwise vorticity dipole (... ~ 1). We theoretically derive a spanwise loading distribution that is parabolic instead of elliptic, and this physically represents the additional circulation around the wing that is associated with reattached flow. This is a fundamental feature of wings with a broad-facing leading edge. The experimental measurements of the spanwise circulation closely approximate a parabolic distribution. The vortex model yields very agreeable comparison with direct measurement of the lift and drag, and the roll moment prediction is acceptable for ... ≤ 1 prior to the roll stall angle and up to side-slip angles of 20#176;. [ABSTRACT FROM AUTHOR]

Delayed detached eddy simulation (DDES) based on Spalart-Allmaras one-equation turbulence model is conducted to investigate the flows over the three-dimensional (3D) S809 airfoil at a wide range of angles of attack (AOA) from 0° to 90°. In addition to 3D DDES, for comparison purpose, simulations of 2D steady Reynolds-Averaged Navier-Stokes (2D RANS), 2D unsteady Reynolds-Averaged Navier-Stokes (2D URANS) and 3D URANS are also performed. The results obtained from 3D DDES have an excellent agreement with the experiment at all the studied AOAs. The 2D URANS and 3D URANS have a similar performance, both overpredicting the lift and drag coefficients at all the separated AOAs. The 2D RANS overpredicts the lift and drag coefficients when AOA is between the stall AOA and about 30°, and underpredicts beyond 30°. In the attached flow regime at low AOAs, all the simulations can give consistent results in agreement with experiment. Visualization of flows shows that 3D DDES can reproduce the realistic 3D flow structures that are incorrectly revealed in the 3D URANS simulations. It is demonstrated that the DDES mode outperforms the RANS/URANS mode in the overall predictions of wind turbine airfoil flows at AOAs from 0° to 90°. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to present the aerodynamic and flow characteristics of a slender body with a 30° swept wing configuration undergoing a limit cycle oscillation using a synchronous measurement and control technique of wing rock/particle image velocimetry/dynamic pressure associated with the time history of the wing rock motion. The experimental investigation was concentrated on 3 main areas: motion characteristics, static and dynamic surface pressures and static and dynamic particle image velocimetry. The tests' results revealed that the lag in asymmetric twin vortices over the forebody switching from the left vortice pattern to the right one exhibits a hysteresis evolvement during the wing rock motion; the asymmetric triple vortices over the forebody interacted with the flowfield over wings appeared to induce the instability and damping moments. The main flow phenomena responsible for wing rock of wing body configuration were completely determined by the forebody vortices. These exhibit apparent dynamic hysteresis in vertical position, which further influences the wing flows, and the dynamic hysteresis of flows yields the damping moments sustaining the oscillations. [ABSTRACT FROM AUTHOR]

The characteristics of the vortex cores over a 20° circular-cone forebody under plasma actuation are studied at 45° angle of attack. Two long strips of SDBD plasma-actuators are installed symmetrically on the plastic frontal cone near the apex. The length of the electrodes is 20 mm along the cone meridian with the leading edge located at 9 mm from the cone apex. Two AC power sources connected to the plasma actuator respectively. The peak-to-peak voltage and frequency are set at Vp-p ≈ 14 kV and f ≈ 8.9 kHz, respectively. The input power for the plasma on is about 19. 3 W. The induced jet flow is pointed at the windward side. Section velocity distribution and pressure distribution are measured in a low-turbulence 3.0m × 1.6m low-speed open-circuit wind tunnel. The Reynolds number based on the base diameter of the circular cone is 5 × 104. The effects of the plasma actuations on the flow physics，such as lateral force，vortex core center position, axial vorticity，vortex core radius，vortex sub-core radius, vortex maximum tangential velocity and circulation are studied. The actuation of plasma actuator on one side pushes the same-side boundary-layer separation point in the windward direction and pulls the other-side separation point in the leeward direction. The suction peak is decreased on the same side and increased on the other side. The separation shear layer and rolled-up vortex core move out on the same side and close in on the other side. In repetition wind tunnel runs of the same model, the flow asymmetry pattern remains unchanged for port-on and starboard-on, while may be changed for plasma-off coses. The plasma actuations tend to increase significantly the size of the vortex core and increase fairly the size of the vortex sub-core, thus，the vortex sub-core becomes more compact relative to the vortex core than that of plasma-off. Vortex core center may be biased from the mid-point between the maximum tangential-velocity points along radial line passing through the vortex core center due to the plasma actuations. The plasma actuation may strengthen vortex maximum axial vorticity，in case the vortex is dispersed for plasma off. [ABSTRACT FROM AUTHOR]

The self-excited oscillation of a wing body induced by fore-body vortices was studied experimentally. To avoid interference, the experimental procedure is divided into two steps: First, the time histories of the self-excited oscillations were obtained using a free-to-roll rig. Then, these time histories were replicated using a forced-to-roll rig driven by a servomotor. Surface pressure and vortex wake are measured simultaneously during the prescribed oscillatory motions. The free-to-roll results show that the tip imperfection significantly influences the rolling motion pattern of a wing body and that the wing body can exhibit limit-cycle oscillations, fixed-point attractors, and apparent chaotic oscillations as the azimuthal angle of the artificial perturbation is varied. The measurements of surface pressures and flow fields indicate that the oscillatory types depend on the switching modes of fore-body vortices with roll angle, and the switching modes are determined by locations of tip perturbations. Furthermore, the fore-body vortices exhibit apparent dynamic hysteresis in vertical position, which further influences the wing flows, and the dynamic hysteresis of flows yields the damping moments sustaining the oscillations. By analyzing the reason for the formation of the rolling moments for sustaining the self-excited oscillations, the flow physics for each type of motion was revealed. [ABSTRACT FROM AUTHOR]

This paper proposes a new tracking control design for agile missiles at high angle of attack regime. Firstly, mathematical descriptions of the problem are presented, taking aerodynamic uncertainties into consideration. Based on the dynamics inverse modeling in a neural network (NN) architecture, the angle of attack command augmented by a variable structure technique is then generated to correspond with the desired turn rate. Subsequently, a nonlinear adaptive control approach using sliding mode control (SMC) technique is employed to track the angle of attack command. Due to the introduction of extended state observer (ESO), the proposed approach is able to online estimate the system uncertainties and calculate derivatives of signals required for composing the control law. Finally, to demonstrate the feasibility and validity of the proposed methodology, the numerical simulation of the agile missile intercepting a target in the rear hemisphere is presented. [ABSTRACT FROM AUTHOR]

To investigate the vortex breakdown over the ONERA70 delta wing at an angle-of-attack of 27°, unsteady simulations were performed using Reynolds-averaged Navier-Stokes and Spalart-Allmaras delayed detached-eddy simulations. A low-diffusive preconditioned Roe scheme with third-order MUSCL interpolation scheme was applied, along with second-order dual-time stepping combined with diagonalized alternating direction implicit method for unsteady simulation. Vortex breakdown was investigated through an examination of total pressure loss, axial velocity, and axial vorticity around the primary vortex. Delayed dtached-eddy simulation provided good agreement with experimental data and predicted all physical phenomena related to vortex breakdown well. [ABSTRACT FROM AUTHOR]

In this paper, the influences of strake wing on unsteady characteristics of aircraft at high angle of attack during rolling oscillation, yawing oscillation, and coupled motion at different frequencies is investigated by using a forced oscillation test rig. The experiments include dynamic force and flow field measurements. As the results show, the rolling moments are affected by the vortices generated by the strake wing. The rolling moment is large and shows strong hysteresis both in the rolling and yawing oscillations. The value of the rolling moment coefficient is small in the coupled motion. The change of the rolling moment is related to the unsteady development of the strake vortices. The characteristics are almost the same for the yawing moment, which may mainly be affected by the sideslip angle. This paper is for the aerodynamic configuration design and the flight control design using flow control method. [ABSTRACT FROM AUTHOR]

A semi-analytical, geometrically-exact, unsteady potential flow model is developed for airfoils undergoing large amplitude maneuvers. For this objective, the classical unsteady theory of Theodorsen is revisited relaxing some of the major assumptions such as (1) flat wake, (2) small angle of attack, (3) small disturbances to the mean flow components, and (4) time-invariant free-stream. The kinematics of the wake vortices is simulated numerically while the wake and bound circulation distribution and, consequently, the associated pressure distribution are determined analytically. The steady and unsteady behaviors of the developed model are validated against experimental and computational results. The model is then used to determine the lift frequency response at different mean angles of attack. Both qualitative and quantitative discrepancies are found between the obtained frequency response and that of Theodorsen at high angles of attack. [ABSTRACT FROM AUTHOR]

With non-zero sideslip angle condition, the changing laws of side force and yawing moment of a slender body at high angles of attack were investigated through wind tunnel testing. An active flow control technique was used to control the slender body's side forces and yawing moments a t non-zero sideslip angle, so as to study the controlling law and the range of effective controlling sideslip angles as well. The results showed that at the condition of α = 55° and the range of sideslip angle β = -24°~ + 24°, the side forces and yawing moments showed linear changing laws at the effective azimuthally angle ranges (± 16°) through changing the equilibrium position of a fast swing micro tip-strake installed on the slender model. When the phantom yaw movement of high agility combat aircraft occurs, this active linear flow control technique could be a new possible flight controlling method for aircraft recovering and safety. [ABSTRACT FROM AUTHOR]

A detailed orthogonal design of a vane-type vortex generator (VG) was performed by varying angles of attack and geometrics including delta, cropped-delta, rectangle and trapezium configurations. Influences of different shapes, angles of attack and geometric parameters of VG on the wake vortex were analysed. Results show that the vortex strength and drag of the triangle VG are minimal while the vortex strength and drag of the rectangular is maximal. The rectangular configuration has the highest wake vortex core orientation, while subsequently followed by cropped-delta, trapezoid and triangle layouts. Due to the fact that mounted height and angle of attack of VG have a significant influence on the wake vortex, a coherent computational campaign was conducted on a Unmanned Aerial Vehicle (UAV) equipped with a VG in front of the topmounted subsonic intake. The aerodynamic interference from VG are numerically examined when the UAV operated at high angles of attack. Research revealed the development of forebody vortex and the effectiveness of flow control by vortex generator on the performance of the intake. Comparing to cases without VG, the distortion index |DC60| of the intake decreased 7.1 % and 5.9 % respectively at the angle of slip ß = 2° and 4°, while the total pressure recovery remains almost the same. [ABSTRACT FROM AUTHOR]

Abstract: A renewed interest in vertical axis wind turbines (VAWTs) has been seen recently. Computational fluid dynamics (CFD) is regarded as a promising technique for aerodynamic studies of VAWTs. In particular, 2D unsteady Reynolds-averaged Navier–Stokes (URANS) is commonly adopted, although past studies on VAWTs revealed the limited accuracy of 2D URANS. This paper investigated the feasibility and accuracy of three different CFD approaches, namely 2D URANS, 2.5D URANS and 2.5D large eddy simulations (LES), in the aerodynamic characterization of straight-bladed VAWT (SBVAWT), with a focus on the capability of the 2.5D LES approach in CFD simulation of high angle of attack (AOA) flow. The 2.5D model differs from a full 3D model in that only a certain length of blades is modeled with periodic boundaries at the two extremities of the domain. The applications of these three approaches were systematically examined in the aerodynamic simulations of a single static airfoil and a 3-blade SBVAWT at different rotating speeds. Their capabilities to predict the aerodynamic forces were evaluated through a comparison with the wind tunnel results obtained by other researchers, with particular attention to high AOA flow beyond stall. Among the three methods, 2.5D LES yielded the best agreement with the experimental results in both cases. Detailed examinations of simulated flow field revealed that 2.5D LES produces more realistic 3D vortex diffusion after flow separation, resulting in more accurate predictions of aerodynamic coefficients in static or dynamic stall situations. It is noteworthy that 2.5D LES cannot capture the effect of tip vortex and vertical flow divergence in VAWTs, which used to be regarded by some researchers as the major cause of overprediction of VAWT power in 2D URANS. In this study, the considerably improved results achieved by 2.5D LES imply that the poor accuracy of URANS method is mainly due to its inherent limitation in vortex modeling. In general, 2.5D LES showed good agreement with experimental results at a relatively low tip speed ratio (TSR), but only fair agreement at a high TSR. Compared with the other two approaches, 2.5D LES is regarded as a more promising and effective CFD tool for investigating the aerodynamic characteristics of VAWTs, particularly their self-starting features corresponding to very low rotation speeds. [Copyright &y& Elsevier]

Abstract: The numerical simulation of the flow for the VFE-2 delta wing configuration with rounded leading edges is presented using the Cobalt Navier–Stokes solver. Cobalt uses a cell-centered unstructured hybrid mesh approach, and several numerical results are presented for the steady RANS equations as well as for the unsteady DES and DDES hybrid approaches. Within this paper the focus is related to the dual primary vortex flow topology, especially the sensitivity of the flow to angle of attack and Reynolds number effects. Reasonable results are obtained with both steady RANS and SA-DDES simulations. The results are compared and verified by experimental data, including surface pressure and pressure sensitive paint results, and recommendations for improving future simulations are made. [Copyright &y& Elsevier]

In view of engineering application, it is practicable to decompose the aerodynamics into three components: the static aerodynamics, the aerodynamic increment due to steady rotations, and the aerodynamic increment due to unsteady separated and vortical flow. The first and the second components can be presented in conventional forms, while the third is described using a one-order differential equation and a radial-basis-function (RBF) network. For an aircraft configuration, the mathematical models of 6-component aerodynamic coefficients are set up from the wind tunnel test data of pitch, yaw, roll, and coupled yawroll large-amplitude oscillations. The flight dynamics of an aircraft is studied by the bifurcation analysis technique in the case of quasi-steady aerodynamics and unsteady aerodynamics, respectively. The results show that: (1) unsteady aerodynamics has no effect upon the existence of trim points, but affects their stability; (2) unsteady aerodynamics has great effects upon the existence, stability, and amplitudes of periodic solutions; and (3) unsteady aerodynamics changes the stable regions of trim points obviously. Furthermore, the dynamic responses of the aircraft to elevator deflections are inspected. It is shown that the unsteady aerodynamics is beneficial to dynamic stability for the present aircraft. Finally, the effects of unsteady aerodynamics on the post-stall maneuverability are analyzed by numerical simulation. [ABSTRACT FROM AUTHOR]

Cai et al.5 developed a vortex stability theory for slender conical bodies and analyzed the stability of vortex pairs over slender conical wing-body combinations under small perturbations. An experimental study is presented in this paper to verify the validity of the theoretical predictions. A sharp-edged flat-plate delta wing is tested in a low-speed wind tunnel. A smoke-laser-sheet visualization technique is used to visualize and measure the positions of the vortex pair, which are found to be symmetric and conical over the wing. The same tests are performed on an identical delta wing model but with a flat-plate dorsal fin mounted vertically in the incidence plane of the wing. Two fin heights are tested. The ratios of the local fin height to the local wing semi-span are 0.75 and 1.50. The test results clearly indicate that the vortices become asymmetric and non-conical over the model with the fin height ratio of 0.75 and recover symmetry and conicity over the model with the fin height ratio of 1.50, providing direct experimental evidence of the theoretical predictions. [ABSTRACT FROM AUTHOR]

Forward blowing from a pair of plasma actuators on the leeward surface and near the apex is used to switch the asymmetric vortex pair over a cone of semi-apex angle 10° at high angles of attack. Wind tunnel pressure measurements show that by appropriate design of the actuators and appropriate choice of the AC voltage and frequency, side forces and yawing moments of opposite signs can be obtained at a given angle of attack by activating one of the plasma actuators. Further work is suggested. [ABSTRACT FROM AUTHOR]

Abstract: A numerical study on flows around a slender body of revolution at a high angle of attack is conducted to investigate the influence of strength and circumferential angle of perturbation on flow asymmetry. A source term model is applied to simulate a real geometrical perturbation near the tip of the slender body. It can greatly facilitate the adjustment of perturbation strength. The results show that the bistable phenomenon does not appear any more at a small perturbation strength. For different perturbation strengths, the energy of the asymmetric perturbation flow experiences a similar spatial exponential growth regime with the same growth rate. The appearance of the bistable phenomenon is closely related to nonlinear saturation of the perturbation flow as perturbation strength increases. [Copyright &y& Elsevier]

Based on the dynamic equation of aircraft flying at high angle of attack, dynamical characteristics of aircraft are qualitatively analyzed, and the flying stability is obtained, which is a function of various parameters. Matching these parameters can yield a stable field of aircraft. [ABSTRACT FROM AUTHOR]

The wind tunnel experiments for high angle of attack aerodynamics were designed from the inspiration of understanding the mechanism and development of an innovative flow control technique. The side force, varying with the different rolling angle, is featured by bi-stable situation, and can be easily switched by a tiny disturbance. A miniature strake is attached to the nose tip of the model. When the strake is stationary, the direction of the side force can be controlled. When the nose tip strake, as an unsteady control means, is swung the flow pattern could be controlled. The results obtained from dynamic measurements of section side force indicate that when the strake swing at lower frequency the side force can follow the cadence of the swinging strake. With increasing frequency, the magnitude of the side force decreases. At still high frequency, the side force diminishes to zero. The side forces could be also changed proportionally. Based on the experimental factors, the mechanism of the asymmetry is discussed. [ABSTRACT FROM AUTHOR]

A numerical investigation of the structure of the vortical flowfield over delta wings at high angles of attack in longitudinal and with small sideslip angle is presented. Three-dimensional Navier-Stokes numerical simulations were carried out to predict the complex leeward-side flowfield characteristics that are dominated by the effect of the breakdown of the leading-edge vortices. The methods that analyze the flowfield structure quantitatively were given by using flowfield data from the computational results. In the region before the vortex breakdown, the vortex axes are approximated as being straight line. As the angle of attack increases, the vortex axes are closer to the root chord, and farther away from the wing surface. Along the vortex axes, as the adverse pressure gradients occur, the axial velocity decreases, that is, λ is negative, so the vortex is unstable, and it is possible to breakdown. The occurrence of the breakdown results in the instability of lateral motion for a delta wing, and the lateral moment diverges after a small perturbation occurs at high angles of attack. However, after a critical angle of attack is reached, the vortices breakdown completely at the wing apex, and the instability resulting from the vortex breakdown disappears. [ABSTRACT FROM AUTHOR]

A flight test research program employing a remotely piloted vehicle (RPV) within high angle of attack range has commenced at the Faculty of Power and Aeronautical Engineering of Warsaw University of Technology, Poland. The initial flights of the scaled model of the "Bielik" aircraft were made with the aim to correlate RPV and full-scale flight stall and departure and spin controllability considering the effects of dynamic scaling laws and Reynolds number. The remotely controlled research aircraft, which is powered by a turbine jet engine and equipped with a flight data recording system, proved to be good source of stability and control data at relatively low cost and without additional risk. The research remotely piloted vehicle and results of its initial flights are presented in this paper. [ABSTRACT FROM AUTHOR]

• The DMDc algorithm is found to have an identification problem. • A variant algorithm is introduced to identify the precise intrinsic dynamics. • Dominant fluid features of a static and an oscillating airfoil are investigated. • The input-output model identified by the new algorithm gives better predictions for different oscillating cases Due to the damage caused by stall flutter, the investigation and modeling of the flow over a wind turbine airfoil at high angles of attack are essential. Dynamic mode decomposition (DMD) and dynamic mode decomposition with control (DMDc) are used to analyze unsteady flow and identify the intrinsic dynamics. The DMDc algorithm is found to have an identification problem when the spatial dimension of the training data is larger than the number of snapshots. IDMDc, a variant algorithm based on reduced dimension data, is introduced to identify the precise intrinsic dynamics. DMD, DMDc and IDMDc are all used to decompose the data for unsteady flow over the S809 airfoil that are obtained by numerical simulations. The DMD results show that the dominant feature of a static airfoil is the adjacent shedding vortices in the wake. For an oscillating airfoil, the DMDc results may fail to consider the effect of the input and have an identification problem. IDMDc can alleviate this problem. The dominant IDMDc modes show that the intrinsic flow for the oscillating case is similar to the unsteady flow over the static airfoil. Moreover, the input–output model identified by IDMDc can give better predictions for different oscillating cases than the identified DMDc model. [ABSTRACT FROM AUTHOR]

This paper presents a novel fixed-time extended state observer (FXTESO)-based fixed-time output feedback control scheme for aircraft with thrust vector at high angle of attack. In order to enhance the robustness of the closed-loop control system, a FXTESO is designed to estimate and compensate model uncertainties, external disturbance and the strong coupling among different channels. Furthermore, theoretical analysis shows that the convergence time of FXTESO is independent of initial estimation errors. For the double integrator systems obtained by FXTESO feedback linearization, a continuous output feedback control is utilized to obtain expected performance and fixed-time stability. The stability analysis for the closed-loop aircraft control system is conducted by Lyapunov theory. The daisy chain method is adopted to realize the control allocation. Finally, several numerical simulations are provided to demonstrate the effectiveness and robustness of the proposed methodology. [ABSTRACT FROM AUTHOR]

An Unsteady Vortex Lattice Method is developed and validated. It is coupled with a decambering methodology to account for viscous effects on the aerodynamic coefficients. Two additional methodologies to select a unique solution when multiple solutions arise have been proposed. The transient nature of the aerodynamic loads of a suddenly moving wing at different angles of attack is examined. Sudden jumps are observed in the CL(t) at post-stall angles of attack. The jumps are followed by the presence of asymmetric solutions, which then decline with time and a change in the solution state. Higher angles of attack see an increasing number of jumps. [ABSTRACT FROM AUTHOR]