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Increasing the pressure in gas turbine cycle is done by Compressor which is one of the most important components of the Gas Turbines. Due to positive pressure gradient, the nature of the flow inside the compressor is complicated and for this reason and because of the duty of the blade for transferring energy to the flow, precise design of the compressor’s blades from the aerodynamic view of point is very important In this study, effects of changes the rotor blade stagger angle on a transonic axial compressor performance curves including efficiency and pressure ratio has been studied. To simulate the complicated three-dimensional flow field in axial compressors, a numerical code is used to solve Reynolds Average Navier-Stokes (RANS) equations. comparison between numerical results and experimental data shows a good agreement. When numerical code are verified. Then the rotor blade twist changes on axial compressor performance have been studied. The results show that the rotor blades twist leads to decrease in compressor efficiency and pressure ratio.

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Rotating stall alleviation in an axial compressor with deployment of air injection at its rotor blade row tip region has been experimentally investigated. Twelve air injectors had been mounted evenly spaced around the compressor casing upstream the rotor blade row. Initially, improvement of the compressor overall performance has been examined through air injection, especially at stall point condition. Instantaneous flow velocities at various radial and circumferential positions were measured simultaneously utilizing hot wire anemometry. These unsteady results, obtained from these latter measurements together with signal frequency analyses, provided to describe the stall inception process and consequent flow induced fluctuations and also alleviation process of stall during the air injection. Results show that a small amount of air injection at the rotor blade tip region can affect the total pressure rise and specifically can increase the compressor stall margin efficiently. Air injection of less than 1% of the compressor main flow rate through the injectors has caused the stall margin to be improved by 9%. Air injection at the blade row tip has caused its beneficial effects to extend throughout the blade whole span, especially while working at the near stall conditions.

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Inlet distortion that may be occurred for various reasons at the entrance of a gas turbine, it is caused to disturbed in compressor performance conditions and also all engine components, so it is very important to investigate its controlling methods. The aim of this paper is numerical simulation of inlet distortion in an axial compressor rotor and active control of the instabilities by the air injection at the blade tip region. Flow simulation of inlet distortion is accomplished at compressor entry with five different geometries of circumferential blockage (amounts of circumferential blockage are: 5%, 10%, 15%, 20% and 25% of the compressor inlet duct). For active control of instabilities, 12 injectors have been mounted upstream of the rotor blade row that distributed in circumferential directions symmetrically. The injection mass flow rate does not exceed 2% of the compressor main flow rate at the design point. ANSYS CFX was used for simulation and the turbulence model of k-ω SST has been used through the calculations. The results show that increasing inlet distortion cause to decrease performance and rotor efficiency. Furthermore, for this rotor modeling condition, in 5% and 10% blockage, air injection can improve the rotor performance, but for more than 10% blockage, a strong wake region is formed after the distortion screen and air injection can cause negative effects on rotor performance. Because the strong instabilities can adversely affect the injectors flow and this method instead of modifying the flow field, make it more non uniform than before.

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Investigation of spike stall formation and its propagation in a low-speed axial-flow compressor is the main aim of this study. Experimental measurements are performed in a low speed axial compressor test rig. Measurement parameters include instantaneous velocity and static pressure at the stall inception process. For this purpose several hot wire probes and a high response pressure transducer is used in data acquisition procedure. Instantaneous fluctuations of velocity at upstream of the blade row show that spike stall inception is accompanied by flow separation from the leading edge of the rotor blade and formation of a vortex subsequently. This vortical structure extends over the blade span. Stall cell propagates with a circumferential speed lower than rotor wheel speed which is equal to 66% of rotational speed in this compressor. Furthermore, wavelet frequency analysis is employed for detail investigation of spike disturbances and capability of this method in distinguishing the spike stall is presented. Wavelet analysis, by representing the temporal variation of frequency spectrum, shows dominant phenomena in the transient process from stable operation to the stall inception condition.

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The numerical simulation of near-stall condition in a passage of an isolated subsonic rotor is studied in detail. The requirements of numerical simulation in order to resolve turbulent spectra around the blade are studied. According to the fact that most of unsteady aerodynamic phenomena incept from blades leading edge, and the role of this part in types and intensity of instabilities, the goal of this paper is to investigate the effects of changes in radius of leading edge of airfoil on flow phenomena in different scales of wave numbers. The governing equations of flow-field are solved using different numerical approaches. Resolution characteristics of different modeling and simulation techniques are investigated. The primary geometry of blade uses a standard NACA-65 series airfoil, which has been tolerated by 50% variation in circular leading edge radius. Mesh requirements of flow simulation for intended purposes are studied in detail and some recommendations are proposed to be implemented in numerical aeroelastic simulations. Accuracy and fidelity of LES results are studied with extraction of power spectra around the blade and the portion of resolved energy is also estimated. Results suggest that the order of accuracy and grid density highly affect the small-scale flow phenomena. The variations in leading edge radius also have great effect on energy distribution among resolved scales.

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In recent years, due to the optimal geometry and lower pressure drop, diffusion control vanes have significant applications, especially in the aviation industry and in subsonic and transsonic conditions. In the current research, the airfoil of the axial stator compressor section designed in the National Aerospace Laboratory of India has been selected as the basic geometry. The goal of optimization is to minimize the pressure drop of the entire fluid flow and consequently reduce the drop rate. The working method in this research is the change in the profile geometry of the blade by changing the parameters of the parsec method, which leads to the creation of new geometries at each stage of the code execution. The used optimization method is developed based on Genetic Algorithm. For the aerodynamic analysis of the generated geometry in each step and extracting the total pressure drop value, the MATLAB code is coupled with Ansys software and in each step, after numerical solution for each generated geometry, the total pressure drop value is extracted and returned to the code. Finally, the work output of the vane is more optimal and with a lower pressure drop, which is finally compared with the original vane and introduced as a suitable alternative. The total pressure drop between inlet and outlet in the optimized vane has decreased by 18% compared to the original vane, and the mass flow rate has also increased by 0.083 kg/s, which is a significant amount. The improvement of various aerodynamic characteristics such as Mach number distribution and pressure and drop coefficients can also be seen between the two basic and optimized blades, which is detailed at the end of the article