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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]

This paper presents a linear compound control strategy for nonlinear aircraft system at high angle of attack. Firstly, extended state observers are designed to overcome model dependence by estimating and compensating the total disturbances including the parameters uncertainties and strong coupling among channels. Secondly, the controlled subsystems are approximately converted to linear integrals based on dynamics compensation linearization, respectively. Subsequently, the generalized predictive controllers are designed based on the integral type prediction model instead of solving the Diophantine equation online, which reduces the computation burden of inverse control incremental matrix. Finally, some simulation results and analysis for aircraft at high angle of attack are carried out to prove the effectiveness of the proposed control strategy. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

The heading reversal of air-to-air missile is designed based on relaxed static stability and aerodynamic torque generated by static instability of air-to-air missile, and the kinematic sign is proposed. The slender body theory and the viscous crossflow theory are used to investigate the aerodynamic characteristics involved in the redirection of air-to-air missile. The slender body is assumed to be incapable of achieving the proposed kinematic sign of completing redirection, and it is necessary to increase control techniques such as aerodynamic rudder surface or direct force. The aerodynamic characteristics of a fin-body configuration are obtained for the angle of attack from 0° to 180°, and the trajectory and attitude of air-to-air missile during the process of heading reversal are calculated. The results show that the heading reversal is finished in 1.15 s, and the process of redirection is safe, quick and stable. Finally, some key technologies are discussed in the process of heading reversal.

Copyright of Aero Weaponry is the property of Aero Weaponry Editorial Office and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Journal of Ordnance Equipment Engineering is the property of Chongqing University of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Wind tunnel experiments are carried out on a 3D wing-in-ground effect at pre- and post-stall angles of attack at a relatively low Reynolds number, Re = 8.8 × 104. The rectangular wing, aspect ratio, AR = 6.4 used in the present work has a symmetric airfoil section, NACA0012 and is studied at different ranges of ground proximity and also compared to a wing of cambered airfoil section, NACA4415 not in ground proximity. Unsteady and time-averaged six-component aerodynamic characteristics, coefficients of Lift, Drag, Side forces, Pitch, Roll, Yaw moments at several angles of attack in the range - 8° < a < 18° are reported. Aerodynamic efficiency and drag polars are studied in lieu of ground proximity at several angles of attack. Effect of ground proximity on the occurrence of stall and CLmax is studied. Numerical calculations using steady-state CFD are also made for quantitative comparison. The effect of ground proximity on span-wise pressure distribution and induced velocity in the vicinity of the wing is also discussed in detail. [ABSTRACT FROM AUTHOR]

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.

Copyright of China Mechanical Engineering is the property of Editorial Board of China Mechanical Engineering and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

To understand the jet flow characteristics of turbofan separate exhaust system, a parametric design method based on the initial Class Shape Transformation function was developed. HBPR and UHBPR turbofan separate exhaust systems were designed. Furthermore, the jet flow characteristics of the HBPR turbofan exhaust system under take-off condition with zero angle of attack were studied based on numerical simulation. The jet flow characteristics of the HBPR and UHBPR turbofan exhaust system under take-off condition with high angle of attack were also simulated. The effects of angle of attack and bypass ratio on the jet flow characteristics were investigated and the related flow mechanisms were analyzed. Results show that the axisymmetric plumes of the HBPR turbofan exhaust system are distributed around the engine axis under take-off condition with zero angle of attack. With the plug wake as the center, the core flow, the fan/core shear layer, the fan flow, the fan/free stream shear layer and the free stream are wrapped around the plug wake from inside out. Vortexes appear in the lee area at the back of the cowl and jet flow under take-off condition with high angle of attack. These vortexes cause cross sectional secondary flow and expose the high-velocity core flow to the low-velocity free stream. The contact area and velocity gradient in the mixing region among the free stream, fan flow and core flow increase. Therefore, the mixture among jet flow and free stream strengthens. So the high-velocity region, the high-vorticity region, and the high turbulence kinetic energy region shorten by 55.1%, 47.7% and 50.9% respectively. The vorticity values and turbulence kinetic energy level peak on the upper side of the exhaust plumes increase by about 30% and 87% respectively. Relative to these parameters from the HBPR turbofan exhaust system, the jet velocity peak value of UHBPR turbofan decreases by 5.5% under take-off condition with high angle of attack. The vorticity values and turbulence kinetic energy level reduce due to decreased velocity gradient in shear layers downstream of the nozzle exit plane. The turbulence kinetic energy level peak on the upper side of the exhaust plumes decreases by 29.3%. The reasons are that the contact area between high-velocity core flow and the free stream decreases due to thicker fan flow and the velocity gradient in the core flow and free stream mixing region decreases because of the lower core flow velocity.

This paper investigates an intelligent method for the motion control of a cable-driven parallel suspension system (CDPSS) in wind tunnel tests, especially at high angles of attack, which is characterized by unsteady and nonlinear aerodynamics. Considering the modeling uncertainties and the complex aerodynamic interference, a composite controller that combines deep deterministic policy gradient (DDPG) and computed-torque is proposed to improve the control performance. The tasks at hand consist in the construction of the training environment based on the dynamic equations and the Markov Decision Process (MDP) design. The supplementary computed-torque control is used to enhance the learning rate of the agent. Then a well-trained agent is applied in the high angles of attack maneuvers control with different examples, including single-DOF and multi-DOF coupled motion. The simulation results show the controller could fulfill the training tasks efficiently, and it turns out to be robust and maintain strong generalization ability despite handling the unlearned tasks.

The landing maneuver of aircraft is considered to be a feasible way to solve the problem of ultra short distance,precise landing of carrier based fixed wing aircraft. The existing researches mainly focus on aerodynamic modeling,nominal landing trajectory optimization and trajectory tracking control in the process of landing,but few on the overall analysis of the landing mission. Based on the dynamic model of aircraft longitudinal landing maneuver,the concept of landing trajectory reachable region is introduced,and the mathematical description of reachable region is given;On this basis,the Gauss pseudo-spectral method with both computational speed and accuracy is used to solve the reachable region problem,which is divided into three trajectory optimization problems. Firstly,the upper boundary of the terminal height of the perching trajectory is solved. Secondly,the nearest trajectory is solved when the terminal height is fixed. Finally,the farthest trajectory is solved when the terminal height is fixed.The simulation results show that:the model and solution method can quickly get the calculation results,the vertical landing trajectory of unpowered aircraft can reach the asymmetric region of narrow upper and wide lower,and the left and right boundary curves have good linearity.

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]

Inspired by birds raising their feathers to control the separation flow, vortex flap has become an approach to control the separation flow over airfoils at high angles of attack.In this paper, the aerodynamic characteristics and physical mechanism of vortex flap controlling separation flow over airfoils under different Reynolds number are numerically studied.The results show that vortex flap is able to greatly improve the lift of airfoil at low Reynolds number under high angle of attack.Its physical mechanism is that the vortex flap makes the core position of main separation vortex closer to the airfoil comparing to the original airfoil, and the vorticity of the vortex center position is also greatly improved so that the low-pressure characteristics caused by the vortex core will affect the flow field on the upper airfoil surface.In addition, the pressure filed on the upper surface of the airfoil is divided by vortex flap into two parts:the low pressure aera in the front and the high pressure aera at the back.However, under high Reynolds number(corresponding to conventional aircraft Reynolds number), the effect of vortex flap on improving the aerodynamic characteristics of airfoil at high angle of attack is far less effective compared to the case of low Reynolds number.This explains why birds can improve lift characteristics through feather lifting, while vortex flap can only be used as drag plates in conventional aircraft.

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]

To understand the jet flow characteristics of turbofan separate exhaust system, a parametric design method based on the initial Class Shape Transformation function was developed. HBPR and UHBPR turbofan separate exhaust systems were designed. Furthermore, the jet flow characteristics of the HBPR turbofan exhaust system under take-off condition with zero angle of attack were studied based on numerical simulation. The jet flow characteristics of the HBPR and UHBPR turbofan exhaust system under take-off condition with high angle of attack were also simulated. The effects of angle of attack and bypass ratio on the jet flow characteristics were investigated and the related flow mechanisms were analyzed. Results show that the axisymmetric plumes of the HBPR turbofan exhaust system are distributed around the engine axis under take-off condition with zero angle of attack. With the plug wake as the center, the core flow, the fan/core shear layer, the fan flow, the fan/free stream shear layer and the free stream are wrapped around the plug wake from inside out. Vortexes appear in the lee area at the back of the cowl and jet flow under take-off condition with high angle of attack. These vortexes cause cross sectional secondary flow and expose the highvelocity core flow to the low-velocity free stream. The contact area and velocity gradient in the mixing region among the free stream, fan flow and core flow increase. Therefore, the mixture among jet flow and free stream strengthens. So the high-velocity region, the high-vorticity region, and the high turbulence kinetic energy region shorten by 55.1%, 47.7% and 50.9% respectively. The vorticity values and turbulence kinetic energy level peak on the upper side of the exhaust plumes increase by about 30% and 87% respectively. Relative to these parameters from the HBPR turbofan exhaust system, the jet velocity peak value of UHBPR turbofan decreases by 5.5% under take-off condition with high angle of attack. The vorticity values and turbulence kinetic energy level reduce due to decreased velocity gradient in shear layers downstream of the nozzle exit plane. The turbulence kinetic energy level peak on the upper side of the exhaust plumes decreases by 29.3%. The reasons are that the contact area between high-velocity core flow and the free stream decreases due to thicker fan flow and the velocity gradient in the core flow and free stream mixing region decreases because of the lower core flow 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 super maneuverability of aircraft is a key factor determining its success or defeat in air combat. The analysis and control of aircraft post stall maneuvering at a high angle of attack can greatly improve the aircraft maneuverability. When the aircraft performs a high-angle-of-attack maneuver, the aircraft's attack angle far exceeds the stall angle; thus, the aerodynamic and aerodynamic moment characteristics are not only strongly nonlinear but also have delay effects and strong coupling characteristics. Moreover, the linear control method based on the linearization of small disturbance hardly satisfies the control requirement because there is no typical leveling state. Traditional nonlinear control methods include nonlinear dynamic inverse, sliding mode control, and robust control for high-angle-of-attack maneuvering of aircraft. However, these methods rely on the accurate model of the aircraft and are greatly affected by modeling errors. To realize the high-angle-of-attack maneuver control for an aircraft with thrust vector, a three-channel decoupling control strategy based on active disturbance rejection control was proposed herein. Based on the public data of the third-generation fighter F16, a thrust vector model was developed. The desired triaxial moments were generated by the thrust vector nozzle combination. Active disturbance rejection controllers were independently designed in longitudinal, lateral, and heading channels. The unmodeled dynamics, uncertainty, and strong coupling between the channels were regarded as total disturbance, which was estimated and compensated online. The angular rate damping feedback term made the closed-loop dynamics of the original aircraft approximate a generalized object, which reduced the design order of the active disturbance rejection controller. As two typical post-stall maneuvers, Cobra maneuver and Herbst maneuver were selected for control strategy verification. The numerical simulation results show that the designed three-channel independent active disturbance rejection controllers can eliminate the strong coupling among channels and realize a high-angle-of attack maneuver for the aircraft with thrust vector. The Monte Carlo simulation results show that the control strategy has good robustness.

Aiming at the difficulty of post-stall maneuvering control modeling and control of advanced aircraft under unsteady aerodynamics, a control method with high control accuracy and fast computation speed is proposed based on Radial Basis Function (RBF) network with minimum parameters learning (MPL) and dynamic surface control (DSC) method. Firstly, the aerodynamic characteristics of post-stall maneuvers are analyzed based on the experimental data of large-scale oscillation wind tunnels, and the key factors affecting the unsteady aerodynamic forces are obtained. Then, an accurate unsteady aerodynamic model is established based on the improved extreme learning machine (ELM) method. Secondly, the influence of unsteady aerodynamic forces on the control of post-stall maneuvers is considered. For the uncertainty of advanced aircraft model, high angle of attack flight control laws based on RBF-DSC are designed. In order to improve the calculation speed of the above control law and optimize the parameters, a post-stall maneuver control law method based on MPL-RBF-DSC is designed, and the stability of the method is proved. The coordinated allocation of the conventional aerodynamic surfaces and thrust vectors is realized based on the daisy chain method. Finally, the typical maneuver simulation of “Cobra” is carried out, which highlights the advantages of the design method in this paper, such as high control accuracy, short calculation time and strong robustness.

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.

This paper reports on flow visualisation and surface pressure measurements over the upper surface of a blunt-edged delta wing model at high angles of attack. The flow structure above the upper surface of the blunt-edged delta wing was found to be different compared to delta wing with sharp leading edge. The flow becomes more complicated especially in the leading edge region of the wing. Currently, there is no data available to verify if the primary vortex could reach the apex of the wing when the angle of attack is further increased. Most prior experiments were performed at the angles of attack, α, below 23° with only a few experiments that had gone to α = 27°. These prior experiments and some CFD works stipulated that the attached flow continue to exist in the apex region of the delta wing even at very high angles of attack above 23°. In order to verify this hypothesis, several experiments at high angles of attack were conducted in Universiti Teknologi Malaysia Low Speed wind Tunnel (UTM-LST), using a specially constructed VFE2 wing model equipped with blunt leading edges. This series of experiments employed two measurement techniques; the first was the long tuft flow visualisation method, followed by surface pressure measurements. The experiments were performed at Reynolds numbers of 1.0x106 and 1.5x106. During these experiments, several interesting flow characteristics were observed at high angles of attack, mainly that the flow became more sensitive to changes in Reynolds number and the angles of attack of the wing. When the Reynolds number increased from 1x106 to 1.5x106, the upstream progression of the initial point of the main vortex was relatively delayed compared to the sharp-edged delta wing. The experiments also showed that the flow continued to be attached in the apex region up to α = 27°. [ABSTRACT FROM AUTHOR]

This paper proposes a finite-time decoupling control strategy for aircraft with thrust vector at high angle of attack maneuver. Firstly, the nonlinear mathematical model of the aircraft is presented. Taking into account the insufficiency of the aerodynamic control surface, a thrust vector model with double nozzles is added. Subsequently, a three-channel decoupling control scheme based on finite-time extended state observer is employed to realize the high angle of attack maneuver. Strong coupling among different channels, aerodynamic uncertainties and other unmodeled dynamics are regarded as total disturbance and estimated by a finite-time extended state observer. Super-twisting (SWT) sliding mode control is utilized to obtain expected performance and finite-time stability. The daisy chain method is adopted to realize the control allocation. Finally, the numerical simulations are provided to demonstrate the effectiveness and robustness of the proposed methodology. [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]

The asymmetric vortices over blunt-nose slender body at high angles of attack result in random side force. In this paper, a nose micro-blowing technology is used to control the asymmetric flow. Pressure measurement and particle image velocimetry (PIV) experiments are conducted in a low-speed wind tunnel to research effects of jet flow rate on asymmetric vortices over blunt-nose slender body. The angle of attack of the model is fixed at 50° and the Reynolds number for the experiments is 1.6×10 5 based on diameter of aftbody. A blow hole (5 mm in diameter) on the nose is processed at circumferential angle θb= 90° and meridian angle γb= 20° with jet momentum ratio Cμ ranging from 5.30×10-7 to 1.19×10−4. Tests are made under two kinds of perturbations. One is called single perturbation with only blow hole and the other is called combined perturbation consists of blow hole and additional granules set on nose. The results show that whether the model has the single perturbation or the combined one, the sectional side force of x/D = 3 varies in the same direction with the increasement of Cμ and remains stable when Cμ is greater than 3.29×10−6. But the stable force values are different according to various perturbations. The fact proves that the size and direction of the side force of blunt-nose slender body can be controlled by the nose micro-blowing.

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 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]

In this study, a model based on potential flow theory is developed to analyze the unsteady nonlinear aerodynamics of a flexible flapping wing at low and high angles of attack. It is assumed that the wing motion is composed of twisting and flapping. The developed unsteady two-dimensional model is applied to the flapping wing using a blade-element-type method. The computational model presented in this study shows better agreement with experimental data especially at a high angle of attack. Analytical results also show that the wings section does not necessarily twist linearly along the wingspan. [ABSTRACT FROM AUTHOR]

The development of a fully reusable vertical-takeoff-and-vertical-landing (VTVL) rocket is indispensable for reducing space transportation costs. However, there are many technical issues associated with such vehicles, such as turnover maneuvers during return flight where the pitching moment plays a key role. It is known that aerodynamic characteristics can be controlled by installing aerodynamic devices, but the relationship between the aerodynamic characteristics and the flowfields has not been explored. To clarify this relationship using computational fluid dynamics (CFD), we investigated these flowfields and aerodynamic characteristics, in the case where we install such devices (fins) in the nose part of a reusable rocket. We found that vortices form downstream of the aerodynamic devices. For angles of attack between 0 and 90 degrees (in which the fins are located in the upstream portion), these vortices significantly affect the surface pressure on the rocket and increase the pitching moment. On the other hand, for AOAs between 90 to 180 degrees (in which the fins are in the downstream portion), the effect of these vortices on the on-surface pressure is negligible, and only vortices formed near the surface of the fins increase the pitching moment. [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 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.

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]

Strong asymmetrical vortices appear on the leeward of slender body at high angles of attack, which has very unfavorable effect on the stability and control of the aircraft. A method is developed to control the side force of slender body at high angles of attack, and is verified in wind tunnel. A thin-film triangular self-excited oscillation flag is fixed at the tip of the slender body model whose semi-apex angle is 10°. Side force is approximately linearly proportional to roll-setting angle of self-excited oscillation flag at high angles of attack, and the slop of fitting straight line obtained by the least square method is −0.158. The linear relationship between side force and roll-setting angle provides convenience for developing side force control law of slender body at high angles of attack. Experimental data shows that the side force coefficients vary linearly with roll-setting angles when a specific plastic self-excited oscillation flag is used as the control flag. The range of side force coefficient and roll-setting angle are, respectively, −3.2 to 3.0 and −20° to 20°. The device is simple, effective, and is of great potential in engineering application.

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]

Investigations have been made adopting experiments and computations on an ogive-nosed slender body at different angles of attack and Reynolds number of 29,000 based on the model base diameter diameter. The results indicated an increase in the side force at large angles of attack, which is mainly due to the presence of asymmetric vortices in the leeward of the body. The inclusion of a rectangular cross-sectioned ring in the initial portion of the body reduced the side force at higher angles of attack. However, significant side force was experienced at lower angles of attack (30° < α < 40°). Use of a ring of 3% height was found to be suitable for reducing the side force at a higher angle of attack. From the results obtained it was observed that a ring if placed at a different axial location alters the flow field and changes the side force at higher angles of attack. Further studies indicated that placing of rings pair at an axial location of 3.5 and 4.5 times the base diameter reduced the side force to a very low value at all the angles of attack for the present shape of body and flow conditions. [ABSTRACT FROM AUTHOR]

Next-generation aircraft and missile are required to have extreme maneuverability, which should maintain stability at high angle of attack. However, the unsteady flow field surrounding the air vehicle would affect the aerodynamics loads and induce unwanted nonlinear motions, of which the rolling motions are usually generated. In order to investigate the unsteady rolling characteristics of a cruciform-finned slender body, the free-to-roll and force measurement tests including particle image velocimetry measurement have been conducted comprehensively. Different types of rolling motion, trimming at equilibrium positions, self-excited rolling oscillation, and self-excited spinning are observed during the free-to-roll experiment depending on the different angles of attack. The force measurement results show that the rolling motions are related to both the static and dynamic stabilities of the rolling moment at balance points. The stabilities of the rolling moment would change as the angle of attack for the model increases. At last, the flow field results from particle image velocimetry measurement indicate that the unsteady rolling motions may be induced by the interaction between the asymmetry vortices and strake wings, fins. [ABSTRACT FROM AUTHOR]

The paper presents the aerodynamic characterization of a low-speed unmanned aerial vehicle, with cropped delta planform and rectangular cross section, at and around high angles-of-attack using flight test methods. Since the linear models used for identification from flight data at low and moderate angles of attack become unsuitable for accurate parameter estimation at high angles of attack, a non-linear aerodynamic model has to be considered. Therefore, the Kirchhoff's flow separation model was used to incorporate the non-linearity in the aerodynamic model in terms of flow separation point and stall characteristic parameters. The Maximum Likelihood (ML) and Neural Gauss-Newton (NGN) methods were used to perform the parameter estimation on one set of low angle-of-attack and one set of near-stall flight data. It is evident from the estimates that the NGN method, which does not involve solving equations of motion, performs on a par with the classical ML method. This may be attributed to the reason that NGN method uses a neural network which has been trained by performing point to point mapping of the measured flight data. This feature of NGN method enhances its application over a wider envelope of high angles of attack flight data. [ABSTRACT FROM PUBLISHER]

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 current research paper is an endeavour to estimate the parameters from near stall flight data of manned and unmanned research flight vehicles using conventional and neural based methods. For an aircraft undergoing stall, the aerodynamic model at these high angles of attack becomes non linear due to the influence of unsteady, transient and flow separation phenomena. In order to address these issues the Kirchhoff's flow separation theory was used to incorporate the nonlinearity in the aerodynamic model in terms of flow separation point and stall characteristic parameters. The classical Maximum Likelihood (MLE) method and Neural Gauss-Newton (NGN) method have been employed to estimate the nonlinear parameters of two manned and one unmanned research aircrafts. The estimated static stall parameter and the break point, for the flight vehicles under consideration, were observed to be consistent from both the methods. Moreover the efficacy of the methods is also evident from the consistent estimates of post stall hysteresis time constant. It can also be inferred that the considered quasi steady model is able to adequately capture the drag and pitching moment coefficients in the post stall regime. The confidence in these estimates have been significantly enhanced with the observed lower values of Cramer-Rao bounds. Further the estimated nonlinear parameters were validated by performing a proof of match exercise for the considered flight vehicles. Interestingly the NGN method, which doesn't involve solving equations of motion, was able to perform on a par with the MLE method. [ABSTRACT FROM AUTHOR]

Computation of flow past high speed vehicles requires the use of a reliable turbulence model. Unfortunately, most of the turbulence models are developed for incompressible flows. Application of these models directly to high speed boundary layers with large density gradients can lead to significant errors in prediction of skin friction. Several compressibility corrections have been suggested in literature to predict these turbulent flows at high Mach numbers. In the present work, we have used two such corrections for the Spalart-Allmaras turbulence model and studied their performance at high angles-of-attack. Flow past an ogive cylinder is considered for the study.

A new dynamic test mechanism has been developed in the unsteady low-speed wind tunnel at Nanjing University of Aeronautics and Astronautics to enhance the abilities of experiments in the unsteady flow cases. The structure of the new test mechanism is described in detail in this paper and the equations used to control the performance of test model’s motions are also listed here. Then a wing-body model was tested in the new test system for the first time, including the forced oscillation tests and free-to-roll test at high angle of attack region. The test results are also discussed. [ABSTRACT FROM AUTHOR]

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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]