مهندس محمود مانی

Attention to plasma actuators as a tool for active flow control has been rapidly growing in recent years. This device, by injecting momentum, accelerates the fluid on the surface, which can be used to control flow separation. In the current study, the control of flow separation in an axial compressor cascade by plasma actuator and boundary layer suction, both separately and simultaneously, has been investigated. To study the effects of the plasma actuator and boundary layer suction on flow control, the flow characteristics, such as the static pressure ratio and the total pressure loss coefficient, have been experimentally evaluated with and without the plasma actuator, as well as the effect of boundary layer suction at different suction rates. Experimental results show that the use of plasma actuator at low speeds reduces total pressure loss and hence improves performance. Moreover, this experimental study indicates that the performance of the compressor cascade with the plasma actuator and boundary layer suction simultaneously, at suction rates of 25, 50, 75, and 100 percent at a speed of 10 meters per second, respectively improves by 13.3, 14, 15, and 17 percent, showing better results compared to using only boundary layer suction.

The DBD Plasma actuator consists of two electrodes which exert the body force resulting from the electrical excitation and is used as a tool for active flow separation control. In the current research, the control of flow separation in a linear cascade of axial compressor blades has been experimentally investigated using a DBD plasma actuator and boundary layer suction. The experimental results show that using the DBD plasma actuator in the chord position of 30%, leads to 31% decrease in the energy dissipation, and also improves the performance of the compressor cascade blades. Likewise, applying the DBD plasma actuator in the velocity of 10 m/s results in the decrement of 19.1% in total pressure loss parameter. Also, utilizing the boundary layer suction device in the same velocity, leads in elimination of low energy part of the flow, and then decreases the total pressure loss parameter up to 14%. The performance of the linear axial compressor cascade blades using a DBD plasma actuator and boundary layer suction has been investigated and reported

Flow control to reduce drag and increase drag and finally increase the ratio of drag to drag (L/D) has always been the focus of aerodynamic scientists. There are many methods to reduce induced drag. The use of DBD plasma operators is one of the newest methods in reducing induced drag. In this research, in order to investigate the performance of DBD plasma actuators, six configurations of plasma actuators have been used as virtual winglets. Experiments have been performed on the wing with the NACA0012 airfoil. These experiments have been carried out in two Reynolds numbers 150,000 and 300,000 and in two voltages, 6 kV and 10 kV and different angles of attack. The results of this research show that the use of this type of plasma actuator at the tip of the wing as a virtual winglet can increase the lift to drag ratio by about 25% in some cases and the use of two small circular and large linear configurations. respectively, they have the best performance compared to other models for use as a virtual winglet.

In this research, complementing previous studies that have focused on flame structure and pollutant levels, the combustion of biogas, composed of methane and carbon dioxide, in a micro gas turbine combustor is experimentally investigated. The influence of the presence of carbon dioxide and variations in its ratio in the fuel mixture on the performance parameters of the combustor and its produced pollutants is examined. The study is conducted at two different thermal powers and constant equivalence ratio for two biogases and compared with natural gas combustion. The results indicate that the carbon dioxide ratio in the fuel significantly affects the turbulent combustion flame structure and performance parameters. Increasing the carbon dioxide ratio in the fuel reduces the combustion rate, causing combustion delay and flame temperature reduction. According to the obtained results, although the addition of carbon dioxide leads to a decrease in the amount of NOx emissions, however, the produced CO in the combustor increases, resulting in an outlet temperature drop and consequently a decrease in the combustor efficiency. Based on these results, it is evident that the utilization of such fuels in combustion systems with swirl burners is feasible. However, measures need to be implemented to mitigate carbon monoxide emissions.

The salient aim of this paper is the shape optimization of an airfoil equipped with Gurney-Flap for aerodynamic performance improvement. The optimization of the slotted Gurney-Flap for improving the aerodynamic efficiency and increasing the lift force of NACA 0012 is the novelty of this research. The Genetic Algorithm, Artificial Neural Network, and Computational Fluid Dynamics are employed for shape optimization. The optimization variables include the height and thickness of the GF; also, the thickness and position of the slot. All analyses have been conducted at Re=0.45×106 and AOA=8o to simulate the take-off phase. After validating numerical results, the optimization process was conducted with two different fitness functions of lift force and aerodynamic efficiency. According to the results, the geometry representing an optimized lift coefficient compared to the geometry with optimized aerodynamic efficiency has a considerably higher height. Furthermore, the thickness of the slot in the first geometry is lower than the second geometry. The first optimized geometry leads to a 21.64 percent lift coefficient increment; there is a 293 percent increment in aerodynamic efficiency due to the second optimized geometry. As a result, it can be indicated that the help of slotted optimized Gurney-Flap can provide the required lift force for Short Take-Off Landing distance.

In this research, the aerodynamic performance of a wing with NASA SC (2) -0714 airfoil and a high-lift configuration, including the main element and the slotted flap at the trailing edge, has been experimentally investigated. Experiments were performed in a wind tunnel at a Reynolds number of 1.01×106 under different angles of attack in the range of -10 to 25° and also flap deflection angles of 0, 20, and 35°. In these experiments, aerodynamic efficiency and wing stall characteristics in different conditions have been determined using force measurements. Also, by measuring the pressure distribution in the middle section of the main element and flap, as well as the total pressure loss in the wake, the onset of the flow separation and the wake profile have been determined, respectively. The results of this study show that by designing the high-lift configuration and using the slotted flap at the trailing edge, the maximum lift coefficient at the flap deflection angle of 35° has increased by 58% compared to the airfoil without high-lift configuration.

In the last decade, the use of plasma actuators for the application of flow control has been very interested. One of plasma actuators types is the Dielectric Barrier discharge (DBD) plasma actuator. Its properties include simple structure, fast response time, low power consumption and lack of moving parts. In this study, the changes of pressure distribution before and after plasma formation at different frequencies and voltages were investigated just above surface at the bottom of the flat plate model and the location of the last plasma actuator. Also the effect of voltage and frequency changes on the pressure distribution has been surveyed. The results of the experimental tests indicate that the plasma produces the ionic jet just above surface, in fact, the plasma actuators produce momentum flow in the boundary layer and produce ionic wind. By increasing the voltage and frequency - more than the breakdown voltage - the momentum flow is created in the direction of the plasma actuator. As the voltage increases and the plasma becomes visible and vortexes are formed on the flat plate.

Effects of flapping on lift and thrust forces in a 3D flapping wing has been investigated at low Reynolds numbers and several reduced frequencies, using experimental tests in a subsonic wind tunnel. Tests has been performed at Reynolds numbers 42000 to 170000 and reduced frequencies 0 to 0.45 that most birds flight at this ranges. Also the ranges of angle of attacks are between 0°-24°. Results has shown that increase of reduced frequency can enhance the lift force by up to 100 percent and in some cases reduce drag force to zero. Furthermore increment of reduced frequency has caused delay at stall of wing. Also by increasing the Reynolds number from 42000 to 86000, major region of the boundary layer of wing surface becomes turbulent, so maximum lift force increases by 40 percent. Wind tunnel test results show that the effect reduced frequency on the lift force was dependent on the angle of attack, so at the lower attack angles, the increase of reduced frequency did not affect the lift coefficient, but, with increment in angle of attack, the positive effect of the reduced frequency on the coefficient of the lift force increased.

In this paper, the development of the DSMC algorithm for flow analysis around axial symmetric complex geometries in rarefied conditions with a computational cost reduction approach compared to a full three-dimensional state, taking into account the appropriate accuracy of the results, as well as the proper selection of particles. One of the challenges is the development of code for complex geometries, as well as the development of code to the functional geometries of axial symmetry. The developed algorithm involves studying different modes of motion and collision of particles with each other or the wall in such a way that the least computations are applied and the high-efficiency solution. Also apply how the extracted condition of axial symmetry is presented for different geometries for each geometry. In the results section, various geometries such as simple geometry (cylinder geometry versus high velocity stream) and complex geometry (including two converging cones with different angles in the airflow of high velocity gas and λ wave formation) were investigated and the results were compared with the results Validated. The results show the accuracy of the proposed algorithm compared with the 3D solvers. Also, the selection of the number of suitable particles is one of the issues that has been studied in order to select the appropriate number of particles. The present results show that the present algorithm is able to use for complex geometries in dilute gas conditions.

In the current research, four different configurations of plasma streamwise vortex generators (PSVGs) for compressible flow control have been experimentally investigated to analyze their capabilities in controlling compressible flow (M=0.428) at the different excitation voltages and frequencies. The impacts of electrical parameters on the performance and efficiency of plasma actuators have been studied. Power spectrum analysis of pressure fluctuations in the boundary layer has been employed to determine the unsteady forcing frequencies of PSVGs. The presence of a dominant frequency in power spectrum diagrams is a strong indication of flow separation in the region from which the pressure signal has been extracted. As such, it was observed that the separation bubble was created in front of the comb-type PSVG when it starts its operation; however, using the T-type configuration diminished the separation bubble. T-type and mesh-type PSVGs, in similar experimental conditions, were observed to be more efficient than the comb-type and saw-type geometries in controlling compressible flow.

Ignition process in a shear-less mixing layer has been studied in this paper. The effect of initial temperature on the flame propagation phase of ignition process is the main goal. The investigation is done by large eddy simulation method coupled with thickened flame approach and DRM_19 chemical mechanism. Mean and RMS axial velocities from both coarse and fine grids and mean mixture fraction have been validated against experimental results. Most upstream and downstream positions of flame edge are in good agreement with experimental. By increasing the initial temperature from 323 K to 1000 K, the mean edge flame propagation velocity increase from 1 to 4.5 m/s. The same trend exists for flame kernel volume. Comparison between calculated edge flame propagation velocity and laminar flame speed and its root density correction show that corrected laminar flame propagation can better predicts the edge flame propagation velocity. Also, by increasing the initial temperature, bibrachial edge flame convert to a triple flame.

In this paper the flow on a finite wing with star damage is numerically and experimentally investigated to understand the influences of damage on the aerodynamic characteristics of wing. To study the effects of different span positions, the damage was considered in tip, middle and root position of the wing span. The wing model of studies is a section of NACA 641-412 asymmetric airfoil with 200mm chord and 800mm span. The aerodynamic coefficients and their increments due to damage were extracted and the results were compared to each other and also to the experimental results. The flow visualization of flow over the damaged wing has been done with pain technique to make evident the flow structure on the model and to understand the influences of damage on the flow. For better understanding of the ability of numerical modeling prediction of aerodynamic performance of a damaged wing, the flow around the wing was numerically analyzed and was validated with the experimental results. A star damage with area about 1% of the wing, can decrease the lift coefficient about 6% and increase the drag coefficient about 15.7% compared with undamaged wing. The star damaged wing also experiences more negative pitching moment coefficient. The results of damage in different locations showed that the damage near to wing tip has less impact on decrement of aerodynamic efficiency.
For better understanding of the ability of numerical modeling prediction of aerodynamic performance of a damaged wing, the flow around the wing was numerically analyzed and was validated with the experimental results. A star damage with area about 1% of the wing, can decrease the lift coefficient about 6% and increase the drag coefficient about 15.7% compared with undamaged wing. The star damaged wing also experiences more negative pitching moment coefficient. The results of damage in different locations showed that the damage near to wing tip has less impact on decrement of aerodynamic efficiency.

Spar ignition in turbulent methane-air jet is studied by a compressible 3D Large Eddy Simulation in OpenFOAM code. The thicened flame model with the one-step chemical mechanism is used for methane combustion. The spar is modeled by artificial enthalpy source in energy equation at the spar location. The validation of the calculations is performed using the experimental results of the turbulent jet of airflow and non reacting mixing of methane jet. The ignition phenomenon and flame propagation are investigated in details for different conditions of initial jet temperature. The results show that the flame propagation speed, the flame lift-off and the flame distance to the axis change with changing the initial temperature.

In this paper, the implementation of the variable-span Morphing wing(VSMW) on the wings of a UAV to achiev the optimal performance for various flight regimes is studied in wind tunnels. For the implementation of this plan, a morphing wing with its related mechanism is designed, built and tested. Tests were carried out at a low- speed wind tunnel with speeds of 35, 60 and 80 m/s in the range of Reynolds Number between 250×〖10〗^3 to 620×〖10〗^3. For the implementation of the morphing wing plan, the maximum deviation of wing span relative to Basewing (BW) is 38 percent and change of aspect ratio is 74% percent. Experimental results showed the improved aerodynamic properties of VSMW compared to BW. These improvements are in the form of reduction of induced drag, thrust required and increasing of lift coefficient, aerodynamic efficiency, endurance and rang. According to the results, endurance and range for VSMW comparison to BW, respectively 17 and 5 percent increased. The experimental studies show that the maximum aerodynamic efficiency of the wing model occurs at 4 degree angle of attack with the maximum value being 14.8 for the base wing, 16.3 for the morphing wing with 50 percent (5 cm) extension in wing span, and 17.4 % for the morphing wing with 100 percent (10 cm) span extension. This research also shows how VSMW can be used as an alternative method of roll control instead of the regular method (i.e. using aileron control surfaces) as it provides more roll control power than the regular method.

In this paper the flow on a finite wing with triangular damage is numerically and experimentally investigated to understand the influences of damage on the aerodynamic characteristics of wing. To study the effects of different span positions, the damage was considered in tip, middle and root position of the wing span. The aerodynamic coefficients and their increments due to damage were extracted and the results were compared to each other and also to the experimental results. Then flow visualizations were practiced to make evident the flow structure on the model and to help to understand the influences of each position of damage on the aerodynamic coefficients. There was the flow through the damage which was driven by the pressure difference between the upper and lower wing surfaces. The flow could take two forms dependent on the angle of attack. The first form was a "weak-jet" which formed an attached wake and resulted in small changes in force and moment coefficients. The second form resulted from increased incidence. This was the "strong-jet" where through flow penetrated into the free stream flow with large separated wake and reverse flow. The effect on the force and moment coefficients was significant in this case. Generally comparing to an undamaged model, increasing incidence for a damaged model resulted increase loss of lift coefficient, increased drag coefficient and more negative pitching moment coefficient.

In this research, experiment and computational fluid dynamics (CFD) are used to assess the performance of UAV with variable-span and sweep morphing wing under low speed conditions. Both wing area and aspect ratio are changed due to variations in span and sweep, whereas structure of the variable-span and sweep morphing wing remains constant. In this study, the numerical results of Fluent software and experimental data are presented. Results are achieved under a low wind speed (50, 60 and 70 m/s). In this case, full extension represents a 30% (10 cm) increase in wing span and 36% (12 deg.) in sweep angle relative to the original wing, with no extension. The results of this study show that the morphing wing is capable to improved aerodynamic efficiency, increased both range and endurance, reduced induced drag and in general reduced thrust required. According to experimental and numerical results, the use of morphing wing can increase the range by 13.6% and 13.5%, also, endurance of the vehicle by approximately 8.855 and 8.17%, respectively. The results of this study show that the maximum value of lift-to-drag ratio occurred at 6 degrees angle of attack and a speed of 70m/s. These results demonstrate that the use of morphing wing improve the lift-to-drag ratio by 10% compared to original wing. Finally, the numerical simulations are compared and show good agreement with the experimental results. This research also showed how morphing concept can be used as an alternative method for roll control.

In the following paper, the accuracy of full turbulent K-e and transitional K-w models in external aerodynamic capturing of NACA0012 and sensitivity analysis have been comprehensively investigated. Among these investigations, some sorts of sensitivity analysis, including changes in free-stream Mach number, free-stream Reynolds number, free-stream angle of attack and generated-grid density have been done. Also the effects of transition inception point and the magnitude of the conformity between numerical results and theoretical concepts about transition and related effective parameters on the magnitude of lift, drag, lift-slope, pressure distribution, and pressure/friction lift-and-drag coefficients have been investigated in detail. To confirm the validity and the accuracy of the results, the numerical results have been compared with some published references.

Boundary layer growth and hence displacement thickness are the key factors in shock position over aerodynamical surfaces at transonic regimes. Since boundary layer growth is a function of free stream Reynolds number and also of the transition location, the free stream Reynolds number and also the other parameters involved in the boundary layer growth and its transition location are determinant in positioning the shock formation and also in aerodynamical forces and moments on a body at transonic regimes. In this paper, the quantitative role of Reynolds number on aerodynamical forces and moments of an AGARD-B model has been investigated. Also, from the results of wind tunnel testing, the blockage of the flow is another effective parameter in the accuracy of the results.