Angle Of Attack Range Research Articles

Numerical Simulation And Aerodynamic Performance Comparison Between Seagull Aerofoil and NACA 4412 Aerofoil under Low-Reynolds

For the aerofoil of small wind turbine working under the conditions of low-Reynolds, the aerodynamic performance of seagull aerofoil and NACA 4412 is contrastively analyzed while the Reynolds-Number is 、 and by the Fluent. The numerical simulation comparison shows that, when the attack-angle range is 0°~20°, the life-coefficient of seagull aerofoil increased 24.97%、46.02% and 48.54% than NACA 4412 respectively; when the attack angle is 4°, the lift-drag ratio of seagull aerofoil increased 40.44%、43.14%和39.40% than NACA 4412 respectively. Under the condition of Reynolds-Number is , the normal working range of attack angle of seagull aerofoil increased 16.7% than NACA 4412 .Under the same simulation condition, the separation point of NACA4412 aerofoil is closer to the leading-edge than seagull aerofoil . The analysis shows that the unique distribution of the thickness and the camber of the seagull aerofoil is the main reason of the better aerodynamic characteristics under the Low-Reynolds.Keywords: Low-Reynolds; wind turbine; aerofoil; numerical simulation; aerodynamic characteristics

Numerical investigation of unsteady aerodynamic characteristics of a pitching missile

Three-dimensional Reynolds averaged Navier–Stokes numerical simulations were carried out to predict the aerodynamic loads of a pitching winged missile based on the finite volume method. The Baldwin–Lomax eddy viscosity model with the modifications suggested by Degani and Schiff was used here. The computational results of the aerodynamic loads of a slender revolution body are also given, and agreed well with the experiment data. The unsteady aerodynamics about the winged missile, which is oscillated in pitching with different frequencies, with different center of rotation and at different range of angle of attack is shown in the paper. It is found that from the computational results, the unsteady hysteresis of the lift, drag and pitch moment coefficients vary not only with the frequency, but also with the position about which the missile performs the oscillatory motions. Although the angle of attack range at which the missile oscillate in pitching are different, the derivative of lift coefficients with respect to the angle of attack is the same for the same phase of oscillating under the same reduced frequency condition.

Aerodynamic forces on fixed and rotating plates

Pressure measurements on static and autorotating flat plates have been recently reported by Lin et al. (2006), Holmes, et al. (2006), and Richards, et al. (2008), amongst others. In general, the variation of the normal force with respect to the angle of attack appears to stall in the mid attack angle range with a large scale separation in the wake. To date however, no surface pressures have been measured on auto-rotating plates that are typical of a certain class of debris. This paper presents the results of an experiment to measure the aerodynamic forces on a flat plate held stationary at different angles to the flow and allowing the plate to auto-rotate. The forces were determined through the measurement of differential pressures on either side of the plate with internally mounted pressure transducers and data logging systems. Results are presented for surface pressure distributions and overall integrated forces and moments on the plates in coefficient form. Computed static force coefficients show the stall effect at the mid range angle of attack and some variation for different Reynolds numbers. Normal forces determined from autorotational experiments are higher than the static values at most pitch angles over a cycle. The resulting moment coefficient does not compare well with current analytical formulations which suggest the existence of a flow mechanism that cannot be completely described through static tests.

New General Approach to Airplane Rotation Analysis

The paper presents a new approach to analysis of dynamic properties of the rotating airplane. The analysis is based on the rate of rotational energy and a full non-linear model of airplane dynamics. It can be applied to a whole range of attack angles. To provide stable rotation, it is sufficient to use algebra and know only the inertial properties of the airplane and the functional relations of either aerodynamic moments or angular accelerations. We prove that the results are fully congruent with those obtained by conventional methods within the scope of validity of the simplified model.

Effect of wing planform on leading-edge vortex structures

Flow visualization experiments are conducted in water tunnel for low aspect ratio cropped wings at low Reynolds number. The experimental results show that the model sweep angle Λ influences the formation and development of the leading-edge vortex. For wings with Λ =0°, the dominant flow structure is transverse vortex. When Λ⩾26°, the dual vortex structure can be observed at some angles of attack, and it is confirmed that the dual vortex is a special structure for flow over low aspect ratio wing at low Reynolds number. For Λ⩾56° wings, the dual vortex structure can be observed in a large range of attack angles. Moreover, in comparison with the outer vortex, the breakdown position of the primary vortex is delayed, and the larger the Λ, the later the breakdown location at the same angle of attack.

INCAS SUBSONIC WIND TUNNEL

The INCAS Subsonic Wind Tunnel is a closed circuit, continuous, atmospheric pressure facility with a maximum speed of 110 m/s. The test section is octagonal ,of 2.5 m wide, 2.0 m high and 4 m long. The tunnel is powered by a 1200 kW, air cooled variable speed DC motor which drives a 12 blade, 3.5 m diameter fan and is equipped with a six component pyramidal type external mechanical balance with a 700 Kgf maximum lift capacity. The angle of attack range is between -45o and +45o while the yaw angle range is between -140o and +216o . The data acquisition system has been modified recently to allow the recording of all test data on a PC type computer using LABVIEW and a PXI – type chassis containing specialized data acquisition modules. The tunnel is equipped with a variable frequency electrical supply system for powered models and a 10 bar compressed air supply for pneumatic flow control applications. In the recent years the subsonic wind tunnel has been intensively used for tests within several European projects (AVERT, CESAR and others). Introduction The 2.5 m x 2.0 m Subsonic Wind Tunnel is the first aerodynamic facility at INCAS. Fig. 1 . Subsonic Wind Tunnel Its main characteristics are the following: • Prandtl type, horizontal single return • Closed test section, atmospheric • Test section size: 2.5 m wide by 2 m high • Speed range: from 7 to 110 m/s • Angle of attack range: -45o to +45o • Yaw angle range: -140o to +216o • Reynolds number (for S l 1 . 0 = ): 1x10 • Contraction area ratio: k = 10 • Turbulence factor: 1.11 The most usual tests involve forces and moments measurements using the external TEM balance, followed by pressure distributions using the existing classical and electronic scanning devices and flow visualizations using smoke, tufts and oil paints. Laser sheet visualizations have been performed recently with spectacular results. Apart from the usual measurements of aerodynamic forces, pressure distributions and visualizations on models of aircraft, missiles, cars, buildings, bridges etc., there have been many special test programs like, for instance wind turbines (tested in the large area settling chamber of the tunnel), sky divers, skiers, cyclists, flag masts etc. The whole experimental activity in the Subsonic Wind Tunnel is carried out according to the existing Manual for Quality Assurance in the Subsonic Wind Tunnel. Description of the facility The tunnel is powered by a 1200 kW, air cooled variable speed (up to 800 rpm) DC motor which drives a 12 blade, 3.5 m diameter fan. The pressure drop across the fan blades is 210 Kgf/m. The motor air cooling is done by means of a special circuit equipped with a 20 KW fan. The DC motor is fed by a thyristor controlled rectifier system connected to a 6 KV, 2 MVA transformer. The tunnel operation, air speed, model positioning and parameter measurements are controlled from an integrated control panel. 12 INCAS BULLETIN No 1/ 2009 DOI: 10.13111/2066-8201.2009.1.1.3

Performance and impact on the seabed of an existing- and an experimental-otterboard: Comparison between model testing and full-scale sea trials

A new door has been designed to reduce hydrodynamic drag coefficient and increase spread of door commonly used in the Mediterranean commercial demersal trawl fisheries. Flume tank testing and engineering sea trials provide data which allow us to illustrate the performance and impact on the seabed of an existing door and a new door design. In the flume tank, each model was tested over a range of attack angles and for a limited range of otterboard heels. Curves of spreading-, drag- and down-force coefficients have been calculated. In the case of sea trials in order to extract the hydrodynamic coefficients an analysis has been applied and a mathematical model was used to calculate attack angle functions. From analysing the differences between engineering sea trials and flume tank tests we have deduced some conclusions about additional ground contact forces on sea trials that affect the performance of the doors. Moreover, a comparison between reaction forces of the flume tank and the estimation of reaction forces at sea has been given. Finally, this study allowed us to notice important differences between traditional and experimental otterboards.

Impulsively generated oscillations in a 3D coronal loop

Aims. The effect of changing the attack angle for the interaction of a fast MHD wave with a 3D coronal loop is studied, to investigate to what extent the properties of the excited transverse kink mode oscillations of the loop depend on this angle. Methods. 3D numerical simulations are performed of the interaction of a fast MHD wave, generated by a pressure pulse, with a 3D coronal loop. The loop itself is modelled as a density enhancement (with a finite plasma beta) within a magnetic arcade. The initial pressure pulse has a width comparable to the loop diameter and is situated outside of the loop, at the same height as the loop apex. This height is kept fixed but the (horizontal) angle between the pressure pulse and the loop is varied. Results. We find that the global, transverse kink mode is efficiently excited for a range of attack angles and qualitatively in agreement with theoretical expectations. The period and damping time are found to be independent of the attack angle. For larger values of the attack angle, the global (longitudinal) slow wave is excited, whereas for intermediate values the second harmonic kink mode is also present.

Variations in Optical Sensor Pressure Measurements due to Temperature in Wind Tunnel Testing

In this paper, wind-tunnel measurements are presented for the airflow fluctuation detection using pressure optical sensors. Twenty-one wind-tunnel test runs for various Mach numbers, angles of attack, and Reynolds numbers were performed in the 6 x 9 ft 2 wind tunnel at the Institute for Aerospace Research at the National Research Council Canada. A rectangular finite aspect ratio half-wing, having a NACA 4415 cross section, was considered with its upper surface instrumented with pressure taps, pressure optical sensors, and one Kulite transducer. The Mach number was varied from 0.1 to 0.3 and the angle of attack range was within -3 to 3 deg. Unsteady pressure signals were recorded and a thorough comparison, in terms of unsteady and mean pressure coefficients, was performed between the measurements from the three sets of pressure transducers. Temperature corrections were considered in the pressure measurements by optical sensors. Comparisons were also performed against theoretical predictions using the XFoil computational fluid dynamics code, and mean errors smaller than 10% were noticed between the measured and the predicted data.

Controlling aerodynamic characteristics of airfoils by means of periodic pulsed energy supply

The possibility of controlling the aerodynamic characteristics of lifting airfoils by means of local periodic pulsed energy supply in transonic flight regimes has been studied. Using a numerical solution of two-dimensional nonstationary equations of gasdynamics, a change in the flow structure near a symmetric airfoil is determined as dependent on the amount of energy supplied to the lower side of the airfoil in a range of attack angles.

Effects of Canard Sweep and Canard-Spanwise Blowing Magnitude on Lift Increment

The effects of different canard's swept angles and jet momentum coefficients on the lift increment of simple close-coupled canard configurations with a 40-deg swept delta wing were investigated in a low-speed wind tunnel by force measurement The experimental results show that at a certain range of attack angles (16-50 deg), on the condition of no blowing, with the increasing of the canard's swept angles, the lift increment and critical stall angles increase, which shows that the canard can be used as a controlling element for aerodynamic forces. On the condition of blowing, when the jet momentum coefficient exceeds a certain value, the more the jet momentum coefficient is, the more the lift increment is, which suggests that indirectly controlling the wing vortex is possible by canard-spanwise blowing for close-coupled canard configurations.

Aerodynamic characteristics of a hybrid trimaran model

Ultra fast marine vehicles can substantially benefit from aerodynamic unloading. The subject of this study is aerodynamics of a hybrid trimaran that comprises three wave-piercing planing hulls and a wing-shaped superstructure. This configuration is characterized by high efficiency and good seaworthiness at speeds about twice those of contemporary fast ferries and combat ships. Aerodynamic coefficients of the above-water structure of the hybrid trimaran were measured in a wind tunnel. A computer code based on the vortex lattice method is applied for modeling steady aerodynamics of this structure in a range of attack angles. The application of an interceptor at the pressure side of the wing is shown to produce significant increase in aerodynamic lift.

511 風車用翼型の性能に及ぼす最大厚み位置の影響(環境とエネルギー)

In this study, the power coefficient was calculated by the blade element-momentum theory for the purpose of bringing out the influence of the maximum thickness position on airfoil characteristics for wind turbine, referring MEL series airfoil data. The power coefficient have different property between Re=1×10^5 and (5.0, 10.0)×10^5. The x_t/C=0.30 case becomes lower attack angle at Re=1×10^5 and every attack angle except (x_t/C=0.30, Re=10×10^5) case stay virtually constant at Re=(5.0, 10.0)×10^5, about the range of attack angle at which the power coefficient exceeds zero. As the maximum thickness position moves forward front, the maximum power coefficient decreases at Re=1×10^5 and increases at Re=(1.0, 5.0)×10^5. Two local maximums of the power coefficient exist with x_t/C=0.42, 0.36 at Re=(5.0, 10.0)×10^5.

Dynamics modeling and performance evaluation of an autonomous underwater vehicle

This paper describes a dynamics model of an autonomous underwater vehicle (AUV). External forces acting on the vehicle, such as hull and control plane hydrodynamic forces, are determined for the full 360° angle of attack range. An accurate through-body thruster model is also incorporated into the simulation. The model is used to simulate two configurations of the C-SCOUT AUV. We use the simulation as a predictive design tool to evaluate and compare the performance of the two configurations before they have been fully developed. Two oceanographic sampling missions are devised to help determine the relative benefits of each configuration.

SCALED HIGH ANGLE RESEARCH VEHICLE SHARV) PROGRAM

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.

Forebody compressibility research of hypersonic vehicle

Three kinds of forebody model of hypersonic vehicles were studied with numerical simulation method. It shows that the two-order compressive ramp model is the best selection among the three for its good evaluative parameters value at the cowl of the inlet. This model can provide higher value of flux coefficient and total pressure recovery coefficient and lower average Mach number compared with those of the other two models. Simultaneously different compressive angles may have different effects. The configuration which the first-order of compressive angle is 4° and the second 5° is the optimum combination. Furthermore factors such as attack angle were concerned. Better result may be obtained with a range of attack angles. Based on the work above the integrated design for forebody/inlet of a hypersonic vehicle was performed. The numerical result shows that this integrated model provides good flow field quality for inlet and engine work.

Flowfield Around Spike-Tipped Bodies for High Attack Angles at Mach 4.5

The requirements for the design of a new short-range high-velocity missile are both the drag reduction and the correct information acquisition for the optoelectronic sensors embedded in the hemispherical nose. High anglesof attack must be studied to fulfill the maneuverability requirements of present and future missiles. A supersonic missile generates a bow shock around its blunt nose, which causes rather high surface pressure and temperature and, as a result, the development of high drag and damage of embedded sensors. The pressure and the temperature on the hemispherical nose surface can be substantially reduced if an oblique shock is generated by a forward-facing spike. Both the experiments and the computations are carried out to study the flowfield around three-dimensional blunt bodies equipped with forward-facing spikes for a large range of attack angles at a Mach number of 4.5. A blunt body, a classical disk-tip spike, a sphere-tip spike, and a biconical-tip spike are studied. The experiments involve high-pressure shock tunnel investigations using a shock tube facility. The differential interferometry technique is applied to visualize the flowfield around the different missile spike geometries. The differential interferogram pictures as well as surface pressure measurements are compared with numerical results. Numerical simulations based on steady-state three-dimensional Navier-Stokes computations are performed to predict the drag, the lift, and the pitching moment for the blunt body and for each spike-tipped missile. The computations allow one to bring out the advantages of each spike geometry in comparison to the blunt body.

논문 : 옆미끄럼이 있는 삼각 날개 / LEX 형상의 와류와 공력 특성

The vortex flow and aerodynamic load characteristics of a sweep delta wing with the leading edge extension in sideslip condition is investigated experimentally. The freestream velocity is 40 m/sec, which corresponds to a Reynolds number per meter of 1.76X based on the wing root chord. The angles of attack range from to 28{^{\circ}}, and the sideslip angles treated are 0{^{\circ}}, -10{^{\circ}}, 20{^{\circ}}. The LEX vortex of the leeward side. The LEX and wing vortics coalesce to to become a concentrated strong vortex or to break down at down at downstream position. Due to the interation of the LEX and wing vortices, a high suction pressure is maintained on the windward wing surface, and a low suction pressure is formed on the leeward wing surface.ﾖ⨀

Experimental study of wind-turbine airfoil aerodynamics in high turbulence

Wind turbines very often have to operate in high turbulence related, for example, with lower layers atmospheric turbulence or wakes of other wind turbines. Most available data on airfoil aerodynamics concerns mainly aeronautical applications, which are characterized by a low level of turbulence (generally less than 1%) and low angles of attack. This paper presents wind tunnel test data for the aerodynamic properties—lift, drag, pitching moment, pressure distributions—of an airfoil used on a wind turbine when subjected to incident flow turbulence levels of 0.5–16% and placed at angles of attack up to 90°. The results show that the aerodynamic behavior of the airfoil can be strongly affected by the turbulence level both qualitatively and quantitatively. This effect is especially evidenced in the angle of attack range corresponding to airfoil stall, as the boundary layer separation point advances along the leeward surface of the airfoil.

Surface pressure model for simple delta wings at high angles of attack

A new aerodynamic modelling approach is proposed for the longitudinal static characteristics of a simple delta wing. It captures the static variation of normal force and pitching moment characteristics throughout the angle of attack range. The pressure model is based on parametrizing the surface pressure distribution on a simple delta wing. The model is then extended to a wing/body combination where body-alone data are also available. The model is shown to be simple and consistent with experimental data. The pressure model can be used as a first approximation for the load estimation on the delta wing at high angles of attack.