جستجوی مقالات مرتبط با کلیدواژه « تونل باد » در نشریات گروه « مکانیک »

Studying the effects of the model scale on wind tunnel tests of aerodynamic vehicles and their components to generalize it to the real scale is a very important because of the direct impact on the performance of the flight system. The aim of the present research is to present a methodology for applying geometric scaling effects of NACA 0012 airfoil the on airfoil aerodynamic performance. AnsysFluent2019R3 software has been used to simulate and investigate the effects of geometric scale on the airfoil aerodynamic performance. Sixteen scale scenarios include changing the Reynolds and the angles of attack assuming Mach constant is 0.256. The airfoil chord length of 50cm (Reynolds 3 million) is considered as the base model in the simulations. The rate of deviation of the validated results for the drag and lift coefficients are 11 and 1 percent, respectively. The results showed that at angle of attack of 10 doubling the airfoil length leads to a decrease of 7.92% in the drag, an increase of 1.25% in the lift and decrease of 18.30% in the pitch moment coefficients. Halving the length of the airfoil at an angle of attack of 15 leads to an increase of 16.29, a decrease of 3.49 and an increase of 9.22% of drag, lift and pitch moment coefficients. One of the important achievements of the present study is the presentation of a methodology for applying the geometric scaling effects in the form of correlations for aerodynamic performance parameters of drag, lift and pitch moment coefficients.

The performance of a Mach 6 Hypersonic nozzle designed to be installed in a reflective shock tunnel has been experimentally investigated in this paper. The purpose of nozzle performance is to create a uniform flow at the outlet and a suitable test time considering the starting flow inside the nozzle. This nozzle is designed using modern optimization methods based on a classic converging-diverging nozzle. Also, the design and construction of a total pressure rake is presented to evaluate the flow quality in the test section. The complexity of investigating hypersonic flows is the sensitivity of the flow to the existence of various types of disturbances and also fluctuations caused by acoustic waves. Therefore, in hypersonic flow, considerations related to the design of the test section as well as the geometry of the pressure rake are very important. In this paper, the dynamics of the reflected shock wave in the shock tube and the starting waves in the test section are investigated. The pressure upstream of the nozzle is in good agreement with the design values with an error of 4%. The Mach number distribution is presented by measuring the pressure distribution at the end of the nozzle. Finally, the Mach number distribution at the nozzle outlet has been compared with the numerical results which shows an error about 3%. The uniformity of the flow at the end of the nozzle, which has been observed numerically and experimentally, shows the effectiveness of the optimal design method of the ultrasonic nozzle.

In the present study the aeroacoustic characteristics of an aerodynamic open circuit wind tunnel is experimentally investigated by furnishing its test section by noise absorbing panels. The objective is to evaluate the effectiveness of sound absorbing walls in reducing the wind tunnel background noise and to investigate the effect of its various parameters such as the thickness of sound absorbing materials on the background noise levels. The results showed that the presence of standing waves across the test section considerably increases the background noise levels for the frequencies higher than about 750 Hz. Using sound absorbing panels at the upstream of the test section improves the acoustic performance of the wind tunnel at mid to high-frequencies while increases the background noise at low-frequencies. The acoustic performance of the wind tunnel is improved at middle frequency range of 150-400 Hz by increasing the sound absorbing materials thickness. The existence of the air gap behind the sound absorbing materials results in a slight improvement in the wind tunnel's aeroacoustic performance at frequencies lower than 300 Hz. Finally, the results showed that using sound transmitting stretched fabric instead of sound absorbing materials results in a significant reduction of the background noise at all frequency range

Wingsuit flying is one of the most popular flight disciplines in recent decades. In the aviation profession, efficiency and safety are paramount concerns for costume designers. An article in this issue examines how waveform changes to the wing surfaces of a wingsuit model improves aerodynamic performance. In order to increase performance, vortices are produced inside the boundary layer that improve the exchange of motion. In this experimental and numerical study, we investigate the formation and evolution of vortices in the Reynolds number range of 106 and provide insights into flow patterns on surfaces with geometric changes. A detailed study of flow structure can be obtained from experimental and numerical evaluations. According to the results, there are significant vortex generators near the backpack due to high pressure. Immediately after the creation of these vortices, the flow is drawn and spread on the surface of the wing in three dimensions. As a result of the angle of attack, the wing surface separates prematurely. Based on the lift and drag coefficients, the study model showed the best performance in flight at an angle of attack of 10 degrees for this flow regime.

Due to the high importance of maneuverability in wingsuit flight, the design and geometry of wingsuits have an important role in improving maneuverability in the flight, therefore active and passive flow control can improve wingsuit performance and maneuverability. One of the ways to improve wingsuit performance is to make geometrical changes such as the structure of the wing surface. In this article, the effects of creating a wave-shaped protrusion on the aerodynamic performance of a wingsuit compared to a smooth wing have been investigated experimentally. By tests in the wind tunnel at different angles of attack for the wingsuit model and measuring the lift and drag coefficients of the model, changing the geometry of the wing surface has improved the performance (L/D) on average by about 20% at angles of attack above 10⸰. Also, by tufts on the surface of the models, the areas of separation were also examined from the perspective of visibility. Experimental observations show that the separation area on the surface of the wavy wing is reduced compared to the smooth wing.

Wingsuit as one of the most popular sports in the field of aviation is very popular. Efficiency and safety are the first priority of the designers of this sport. In this article, the modeling, formation, and evaluation of the surface of the wingsuit during a flight are discussed. Since the wingsuit is under air pressure inside it, the formation of the structure of the surface of the suit changes according to the flight conditions. The surface of the wingsuit is important for the aerodynamic evaluation of the wingsuit model. According to the way of sewing wingsuit clothes, the wave structure will be obtained on the surface of the wing. In order to better observe the effects of changing the wing geometry on maneuverability, the model has been selected as rigid. The experimental results on the surface of the model show that with the increase of the angle of attack, the flow on the surface of the wing increases the disturbance in the middle area of the wing and reduces its effects in the upper area of the wing. This interaction of the flow structure on the model leads to the different performances of the model in different angles of attack; By measuring the forces on the model, AOA= 15 was obtained as the best angle of flight performance for the wingsuit model.

In general, rotating objects always produce different rotational torques according to different dimensions and rotational speeds. In some cases, it can cause a lot of damage to equipment, so it is necessary to be aware of the amount of torque that rotating objects produce at different rotational speeds, as well as in the presence or absence of air flow. in this study, numerical and experimental analysis of non-continuous flow around a cylindrical model with vertical plates under forced rotation is performed and its main purpose is to measure the torque of rotating objects in the presence of wind current also to stabilize rotation. Rotational speeds are constant. First, an aerodynamic torque measuring device was built and then a 3-fin cylindrical model was tested in a wind tunnel, and numerical simulations of 3 and 2 fin models were performed in the same laboratory conditions by Ensys Fluent software. A good agreement was observed between the experimental and numerical results and the maximum error between them was less than 10%, which is acceptable. From the simulation results, it was observed that in every 180 degrees of rotation that the maximum cross section of the models is exposed to direct wind flow, the maximum torque produced by the 2-blade cylindrical model is 30% higher than the maximum torque produced by the 3-blade model. As the wind speed increases from 20 to 60 meters per second, the torque of the 3-fin model increases from 0.4 to 1.2 Nm, which is equivalent to 200%.

Today, measuring the flow parameters of the oscillating airfoil wake is of great importance due to its wide applications in industry and nature. Therefore, the aim of this research is to provide an innovative method for the experimental measurement of the NACA0012 oscillating airfoil. For this purpose, three one-dimensional hot-wire sensors were used, one of which is fixed, and the others are mobile. By using MATLAB code and the connection between the data of these three sensors, the spatial and time correlation of the flow velocity at all points of the wake was obtained. The results showed that measuring different parameters of the wake with less hardware equipment than other researches is possible. In this method, the flow parameters have been measured with higher accuracy by eliminating the errors caused by the large number of sensors and by increasing the number of data collection points in the wake. The results showed that the combination of the free stream flow momentum with the oscillating momentum of the airfoil creates a sinusoidal flow with the same frequency of the airfoil oscillation in the wake. Moving vortices area and the area of diffusing vortices momentum form the two parts of the oscillating airfoil wake.

When a flying vehicle approaches a surface of water or land, changes occur in the pattern of the fluid flow field around it. This change in flow field eliminates the direct effect on aerodynamics and control of the vehicle. This is more common when the vehicle is landing and taking off, as well as flying at low altitudes, which is called the surface effect. In this research, the phenomenon of surface effect and its effect on aerodynamic coefficients and flow pattern around NACA0012 airfoil in the static incompressible subsonic regime have been investigated numerically and experimentally. Experimental tests were performed in the incompressible subsonic wind tunnel of the Ghadr National Aerodynamics Research Center of Imam Hossein University with a cross-sectional area of 80 by 100 cm. The simulation of the phenomenon is a fixed ground with the minimum possible thickness of the boundary layer in the wind tunnel. Solve the flow field numerically based on Navier Stokes equations along with the Transition-SST viscous model. The impact of the surface effect phenomenon on the change of aerodynamic coefficients has been investigated by considering different distances from the surface in the static state. The pressure distribution on the airfoil surface is measured by an accurate pressure sensor and is due to the surface effect phenomenon at close distances to the surface. The results of the static analysis show an increase in lift force and a decrease in drag force.

In this research, the turbulent isotropic flow has been experimentally investigated. Hence, two different grids are made and a contraction channel is installed behind it inside the subsonic wind tunnel to generate an isotropic turbulence flow.  The grids with mesh sizes of 2/54 cm and 5/08 cm were cut on the wood with obstruction ratio of 0/34 and 0/17, respectively. One-dimensional hot wire was used to determine the perturbation velocities in the direction of flow, and an approximation was used to determine the components of other directions. At speeds of 5 m/s and 10 m/s, experiments were performed for each of the grids, which range from a Reynolds number of 8500 to about 33000. To determine the onset of the isotropic location, methods of velocity skewness, kurtosis, turbulence intensity, dissipation rate, and longitudinal scales such as Kolmogorov and Taylor lengths were used. For skewness and kurtosis, the numbers show 0 and 3, respectively, which indicate the isotropic flow. Results showed that with increasing the velocity, the isotropy of the flow was delayed. Also, in a grid with a lower obstruction ratio, the intensity of turbulence will be less near the grid, but as it moves away from the grid, the intensity of turbulence will increase.

In this paper, the characteristics of the ventilated cavitation around an axi-symmetric body with a cone nose at different water velocities and injection rates have been investigated experimentally. The experiments are done in an open loop water tunnel with a maximum water flow of 40 m/s. The main purpose of this research is to study the rate of air injection and the flow velocity on the coefficient of drag and the relevant super-cavity shape. The effect of air injection rates from end of the cone nose of the body on two phases fluid flow characteristics was studied for different water velocities within the tunnel. For each case of input conditions, the variations of the drag force and pressure distribution profiles along the length of the body and the length of the super-cavity are measured. The results of the present study show that increasing of the injection rate to an optimal value has a significant effect on the reduction of the drag force and after that the reduction of the drag force is negligible.

One of the important issues in civil engineering is creating a calm place and a sense of security for the residents of high-rise structures against the force of earthquakes and wind. Therefore, the use of control systems has been considered under dynamic loads. Tuned Liquid Damper is Affordable and useful device for controlling the vibrations of structure under dynamic lateral loads. In this study, the standard high-rise structure has been modeled in ANSYS software under earthquakes (far and near-field) and wind profile with a speed of 100 meters per second. The wind tunnel has been simulated and the interaction between wind and structure has been investigated. In order to reduce the responses of high-rise structures under far-field records (Elcentro 1940 and Hachinohe 1968) and near-field records (Northridge 1994 and Kobe 1995) and wind force, a TLD with a cube tank was used. The responses of the structure such as displacement, velocity, acceleration, pressure over to the structure and wind streamline over the walls of the structure have been analyzed and also, the aerodynamic and aeroelastic behavior of the high-rise structure against wind has been investigated. Averagely, the simulation results show that the Tuned Liquid Damper could reduce the maximum displacement of the structure to 16% under far-field records, 0.5% under near-field records and 13% under wind force.

In the present study, the wake flow field of a submarine model was investigated experimentally in a wind tunnel. The experiments were conducted to determine the effect of the location of control surfaces on the wake inflow to the impeller of the submarine. In order to investigate the effect of the location of control surfaces as the most important innovation of the present study, the aforementioned surfaces were installed in three longitudinal positions X/L=0.89, 0.92, 0.95 on the heel of the submarine model, and the wake flow was measured at position X/L=1.7 and the Reynolds number 6*10^5  by a five-hole probe and a hotwire anemometer. Finally, the longitudinal position X/L=0.95 was selected as the optimal location for the stern planes to improve the wake inflow to the impeller in terms of reducing its total area and the least amount of turbulence and non-uniformity. The results obtained during this study showed that arriving of the holder basechr(chr('39')39chr('39'))s wake to the stern area increases ​​the area and average velocity and subsequently reducing the non-uniformity of the wake flow.

A new experimental technique has been developed to measure the pressure distribution over the surface of a spinning wind tunnel model. The technique is unique in that all elements of the instrumentation, thus avoiding many of the technical problems and operational limitations associated with previous attempts to measure this effect. Surface pressure distributions were obtained for selected tip speed ratios for different angles of attack (5000rpm). The results obtained from the pressure profiles determine the Magnus forces and make it possible to interpret the boundary layer and the effects of separation. In this method, in addition to determining the Magnet force, its distribution on the model is also obtained. The results show that most of the Magnus force is created at the ends of the projectile. The validity of the data was established by comparing the integrated pressure values with directly measured balance data. Similar results were obtained by the numerical simulations and were compared with the experimental data. This new technique can be applied to a variety of model configurations and Mach number regimes.

Galloping is a large-amplitude, low frequency, wind-induced oscillation of overhead power transmission lines with one or multi loops of standing waves per span which occurs due to wind flow. Based on the field data, numerous galloping oscillations occurs in the form of one loop oscillation which whereby high dynamic loads are imported to the support structures. In this research, the results of wind tunnel tests have been performed on a two-span distorted scale model with an ice-accreted cross-section under uniform and non-uniform aerodynamic loadings. Dead-end and suspension insulators have been applied to the support points. Then, based on identifying the most critical state of the lines oscillation, a solution has been proposed based on increasing their bending strength through the application of hardening local covering. The results showed that the most critical state of the cables oscillation in the galloping is related to the one-loop oscillation, which occurs as a result of interactions between the cables of adjacent spans under uneven aerodynamic loading and the use of suspended insulators, and the dynamic forces applied to the supports are about 20% more than the case when the cables oscillate due to the dead-end connections attached to the support structure. Also, applying the local covering with a length of 20% of cable span leads to a 27% reduction in dynamic support reaction of one-loop galloping.

The purpose of this paper is to investigate and compare the aerodynamic coefficients obtained from the wind tunnel, numerical solution (Fluent) and engineering software (MD) for a cruise missile. The results are obtained in zero deflection of the control surfaces. For this purpose, the analysis has been carried out on the aerodynamic coefficients of the three Mach numbers: 0.6, 0.75, and 0.85, and various angles of attacks. The results of the numerical solution for calculating the coefficients of the lift, drag, normal and axial forces are respectively with a mean difference of 8.6, 1.7, 8.3 and 8.4 percent, respectively, in comparison with the wind tunnel. The results of the MD software for drag and axial forces are acceptable with an average error of 11% and 20%, respectively. Also, the existence of errors in the MD software, such as taking into account the effects of the air inlet opening only in the axial direction, shows that this method is unreliable in the present study. The results show that there is a great similarity between the behavior of the aerodynamic coefficients changes relative to the angle of attack in all three experimental and numerical methods and the MD software. Also, the pitching moment coefficient variation according to the angle of attack indicates that the trim angle varies from +6 to + 7 degrees.

Aerodynamic characteristics of an aircraft with delta wing are considerably affected under the ground effect in take-off and landing phases. In this research, static ground effect of a 60 degrees delta wing with axisymmetric sharp edges in combination of body and vertical tail at low speed wind tunnel are investigated. The wind tunnel is closed type has an opened test section that its dimensions is 2.8 m × 2.2 m and maximum velocity is 90 m/s. In these tests ground effect is simulated using a fixed plane that its height is variable. With decreasing the height from the ground plane, the lift force is increased, induced drag force is decreased, total drag force is increased and nose-down pitching moment is increased nonlinearly. The rate of increasing of lift coefficient in linear regions increased with decreasing of height from ground plane. At the positive angle of attack the most percentage increasing of lift coefficient is due to 5 degrees angle of attack as the result of existence of perfect vortex flow over the whole wing surface. The minimum case is due to 30 degrees of angle of attack.

One of the conventional ways to increase velocity of a fluid flow is using a nozzle in stream wise. In this paper, in order to increase the flow velocity in a low speed wind tunnel, and to carry out experiments at higher velocities. An optimal nozzle profile, by considering constraints such as the nozzle inlet cross section, inlet velocity, and nozzle length has been designed using computational fluid dynamics. Design variables were the nozzle inlet cross section, outlet velocity, the turning point of the nozzle profile, the uniformity of the outlet flow and the effective length of the test section. The performance of the designed nozzle was experimentally investigated. According to the results, the velocity increased by 13 m/s and the uniformity of the flow decreased by ±1%.

In this paper, the design of a diffuser for a suction type subsonic wind tunnel equipped with three axial fans has been taken into account by CFD. The tunnel test section has a 140 cm height and 195cm width. For numerical modeling the axial fans that contains 10 rotor blades and 13 stator blades 3D modeling to simulate the impeller have been used. The 3D modeling to simulate the impeller method provides an adequate means for simulating the swirl effects of air flow. The best geometrical characteristics have been selected for designing the optimal performance for the diffuser. The final design includes a circular diffuser which incorporates a 4 degree equivalent cone angle. This study also suggests that the circular scheme is superior to the triangular and fitted design aspect of the total pressure loss and the volumetric flow rate.

In this paper, the inception, growth and development of cavitation are investigated around the combined body experimentally. The diameter of afterbody and the length of model are 25 and 210 mm respectively. The angles of conic nose are 〖30〗^° and 〖45〗^°. Models were tested in a high speed closed circuit cavitation tunnel. A peripheral groove is established on the afterbody and then one conic nose, experiments are repeated in these cases. A comparison was done between all cases. For 〖30〗^° model with groove on afrebody and without groove, cavitation is initiated at a small distance behind the body. If the groove is established on the nose then cavitation is initiated into it. In 〖45〗^° model, cavitation is initiated on the interface of the nose and afterbody. Into the groove and behind of afterbody, cavitation will be initiated at the center of vortex resulted by separation. By developing the length of cavitation on the afterbody, regular longitudinal oscillations will be occurred. Behind the model, inside the wake of body, cavitation initiates at the center of annular vortex. After developing the length of cavitation area ,bubble shedding occurs at the end of cavitation region. During the fluctuating of cavitation area, bubble shedding occurs randomly. In these cases, intense noise is heard.