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To extract power from the wind, the high enough velocity of the wind is a key factor. However, winds with high enough velocities are not available in most areas; therefore, finding a method to enhance the velocity potential of the low-speed winds remains a serious challenge in wind engineering. In this study, a hollow duct with a specific configuration, called Multi-Stage Inducing Duct, is presented which is capable of significantly enhancing the velocity profile of low-speed winds, enough to start most of the small horizontal wind turbines. For example, a three-stage inducing duct can increase the wind velocity from 3 m/s to 9.3 m/s on average at the first-stage throat of the duct, which equals the increase rate of 29 times in wind potential power. This configuration of the duct was introduced and numerically simulated. Moreover, the effects of its different geometrical parameters were investigated. The impact of wind velocity and the number of the stages on the duct performance were investigated. Finally, a study of the relative effectiveness of adding each of the stages was carried out to help decide where the upper limit for the number of the stages is.

This paper describes the experimental study of surface pressure over a supercritical airfoil which was oscillated in pure pitching, pure plunging and combined pitch-plunge motions at the Reynolds number of 8.76*105. While the surface pressure distribution is of significant importance in stability and performance of an airfoil, not sufficient information is available on the pressure distribution in dynamic stall. The experiments were conducted in a closed-loop wind tunnel utilizing pressure transducers array. The motions were designed to maintain constant reduced frequency, Strouhal number and phase difference. Three different regions were assumed to represent the pressure distribution over the airfoil. The results showed that LEV formed on the upper surface manifested different behavior. In the attached flow region the LEV grew and shrunk over the upper surface but in the light stall region the LEV spilled on the airfoil while a small partial LEV remained at the leading edge. In the deep stall region the LEV spilled entirely and the flow was fully separated. The formation of Laminar Separation Bubbles and suction peaks were also reported in low angles of attack. Besides, the pitching moment Damping Factor was studied to determine the level of airfoil stall flutter stability. For lower amplitudes of pitching motion the airfoil seemed to be stable except where deep stall occurred. However for high amplitudes the airfoil had a tendency to enter the stall flutter. Nevertheless, forcing the airfoil to undergo a combined motion improved the stability condition in all cases.

This study provides a direct method based on the method of characteristics with assumptions of non-viscous flow and thermally perfect gas to design an axisymmetric plug nozzle. Inputs for numeric code include output Mach number, specific heat ratio, global gas constant, and a number of discrete steps of Prandtl-Meyer expansion fan. In the design process, no simplification was made and only unique spike geometry was created. Based on the criterion that locates the tip of the spike on the axis of symmetry, for each output Mach number, a correction factor was extracted which is the sole source of theoretical errors. The studied parameters include spike geometry, parallelism of exhaust nozzle flows and uniformity. Based on the numerical analysis, obtained results are consistent with desired output Mach number, and output flow is parallel to the axis of symmetry. A simple and direct formulation was used which performed very fast. Due to the error detected, which resulted in the surface slope reduction factor, some movement in spike geometry relative to ideal conditions was raised. Thus, the effect of a reduction factor may tend to zero by increasing the design Mach number.

Experimental results of surface pressure distribution over a thin supercritical airfoil and its wake are presented. All tests were conducted at free stream Mach numbers ranging from 0.27 to 0.85 and at different angles of attacks in a transonic wind tunnel. The model was equipped with static pressure orifices connected to high-frequency pressure transducers. The present paper evaluates variations of shock wave location with both Mach number and angle of attack variation, as well as its interaction with the boundary layer, leading to the buffet phenomenon. Note that, for this thin supercritical airfoil, there exist only a few experimental results regarding surface pressure distributions, corresponding forces and moments, and the shock wave oscillations and its behavior with various flow conditions. The frequency of the shock wave oscillation and unsteady wake behavior at a freestream Mach no. of = 0.66 and at different angles of attacks are measured by the cross-correlation technique by means of pressure sensors located on the suction side of the model and via the rake total pressure data that was traversed vertically behind the model, respectively. From the analysis of surface pressure distribution and wake data, drag divergence occurred at a certain angle of attack and at a frequency equal to the shock wave oscillation frequency.

A series of experiments were conducted to study impacts of the free-stream Mach number, back pressure, and bleed on the stability of a supersonic intake. The flow stability is related to the buzz phenomenon, i.e. the oscillation of all shock waves of the intake, which may also occur when the intake mass flow rate is decreasing. In this study, the intake was axisymmetric with Mach number of 2.0. The results showed that stability margin of the intake decreased when the freestream Mach number increased for both bleedoff and bleed-on cases. In the configuration without bleed, the frequency of buzz oscillation increased when the freestream Mach number decreased or when back pressure increased. By applying bleed and, consequently, preventing separation of the flow, the intake became more stable and the shocks oscillated with a smaller amplitude during the buzz phenomenon. Also, when the bleed was applied, the buzz triggering mechanism varied from the Dailey criterion to the Ferri one, which considerably changed stability characteristics of the intake.

Purpose This paper aims to analyze the influence of the changings in geometrical parameters on the aerodynamic performance of the control canard projectiles. Design/methodology/approach Because of the mentioned point, the range of projectiles increment has a considerable importance, and the design algorithm of a control canard projectile was first written. Then, were studied the effects of canard geometric parameters such as aspect ratio, taper ratio and deflectable nose on lift to drag coefficient ratio, static margin based on the slender body theory and cross section flow. Findings The code results show that aspect ratio increment, results in an increase in lift-to-drag ratio of the missile, but increase in canard taper ratio results in increasing of lift-to-drag ratio at 1° angle of attack, while during increasing the canard taper ratio up to 0.67 at 4° angle of attack, lift to drag first reaches to maximum and then decreases. Also, static margin decreases with canard taper ratio and aspect ratio increment. The developed results for this type of missile were compared with same experimental and computational fluid dynamic (CFD) results and appreciated agreement with other results at angles of attack between 0° and 6°. Practical implications To design a control canard missile, the effect of each geometric parameter of canard needs to be estimated. For this purpose, the suitable algorithm is used. In this paper, the effects of canard geometric parameters, such as aspect ratio, taper ratio and deflectable nose on lift-to-drag coefficient ratio and static margin, were studied with help of the slender body theory and cross-section flow. Originality/value The contribution of this paper is to predict the aerodynamic characteristics for the control canard missile. In this study, the effect of the design parameter on aerodynamic characteristics can be estimated, and the effect of geometrical characteristics has been analyzed with a suitable algorithm. Also, the best lift-to-drag coefficient for the NASA Tandem Control Missile at Mach 1.75 was selected at various angles of attack. The developed results for this type of missile were compared with same experimental and CFD results.

The effect of porosity in oscillating situations (to the authors' knowledge, for the first time) on a supercritical airfoil (SC0410) has been experimentally investigated. Tests have been carried out in an open circuit suction-type wind tunnel at a free stream Mach number of M = 0.80. Both static and dynamic (pitching) tests have been carried out on the mentioned airfoil. The oscillation frequency for the unsteady tests has been set to 3 and 6 Hz. The amplitude of frequency is ±1 deg. The effect of porosity has been surveyed on the magnitude of pressure fluctuations, phase shift, and lift coefficient loop. The investigations show that increasing porosity in the test section of transonic regime, contrary to the impression, does not necessarily improve results.

Wind tunnel experiments were conducted to evaluate surface pressure distribution over a semi span swept wing with a sweep angle of 33 degrees. The wing section has a laminar flow airfoil similar to that of the NACA 6-series. The tests were conducted at speeds ranging from 50 m/s to 70 m/s with and without surface roughness. Surface static pressure was measured on the wing upper surface at three different chordwise rows located at the inboard, middle, and outboard stations. The differences between pressure distributions on the three sections of the wing were studied and the experimental results showed that roughness elements do not influence the pressure distribution significantly, except at the inboard station. The spectral analysis of the pressure-time signals acquired from the pressure orifices over the wing upper surface showed that roughness had significantly affected the zero frequency amplitude. In this study, the zero frequency amplitude and its variations with roughness elements was investigated at three different chordwise positions.

A series of wind tunnel tests are performed to examine the flow field over a swept wing under various conditions. The wing has a laminar flow airfoil section, similar to those of the NACA 6-series. Static pressure distributions over the upper surface of the wing, in both chordwise and spanwise directions, are measured at different angles of attack. The data are employed to predict the transition point at each chordwise section. The skewness parameter of the pressure data shows that this factor drops to zero in the transition region. A comparison of the calculated transition point on the wing surface with that obtained from the 2D computational method shows reasonable agreement over a portion of the model. The power spectral density calculated from the total pressure data of the boundary layer, over the wing surface, at several locations shows the instability modes.