In this paper, a parametric study of COMPRESSOR performances is performed by the streamline curvature method. Effects of three input parameters in the design process, e. g., number of blades, distribution of blade thickness, and blade sweep angels, on the main objective parameters in the aerodynamic design, e. g., velocity distribution, efficiency, and pressure ratio, are investigated in the parametric study. Initially, a certain two-stage AXIAL COMPRESSOR is designed by the streamline curvature method. Validation of the results is confirmed by comparing the obtained results with the experimental ones. Regarding various values for the aforementioned input parameters, the first stage of the AXIAL COMPRESSOR is redesigned, and the output parameter is established. Therefore, the sensitivity of the design results to each of the aforementioned parameters is recognized. Results show that increasing the blades sweep angle causes the flow behavior, such as efficiency and pressure ratio in the AXIAL fan, to improve while reducing it provides a completely contrary result. Also, reducing the rotors blades number leads to an increase in the pressure ratio and efficiency while its increase causes a contrary result. It is concluded that a reduction in the number of the blades has a stronger effect on the performance parameters than when it increases. The results also show that the effect of the thickness in the hub is greater than the thickness of the tip, and its increase leads to reduce both efficiency and pressure ratio.

Improving the efficiency of COMPRESSORs has been one of the most important goals of researchers over the years. In this paper, three different methods are presented for parameterization and blade optimization of AXIAL flow COMPRESSOR. All methods consist of flow analysis tool, optimization algorithms, and parametric geometry generation tool that are different in each approach. An objective function is defined based on the aerodynamic performance of blade in the acceptable incidence angles range. A double circular arc blade is used as the initial guess for all methods. The performance of optimized blades and the initial blade are compared to evaluate the capability of various methods, and a good agreement is achieved. The results show that the level of performance improvement in each method depends on the number and type of the chosen parameters. All three methods have improved blade performance at the design incidence angle. However, only the first method shows significant performance improvement in off-design conditions.

Design models of multi-stage, AXIAL flow COMPRESSOR are developed for gas turbine engines. AXIAL flow COMPRESSOR is one of the most important parts of gas turbine units.Therefore, its design and performance prediction are very important. One-dimensional modeling is a simple, fast and accurate method for performance prediction of any type of COMPRESSORs with different geometries. In this approach, inlet flow conditions and COMPRESSOR geometry are identified and by considering various COMPRESSOR losses, velocity triangles at rotor, and stator inlets and outlets are determined, and then COMPRESSOR performance characteristics are predicted.Numerous models have been developed theoretically and experimentally for estimating various types of COMPRESSOR losses. In the present work, performance characteristics of the AXIAL-flow COMPRESSOR are predicted based on one-dimensional modeling approach. Firstly, the proposed algorithm for modeling and then the losses model for calculation of pressure loss coefficient in the blades cascade have been represented. In this study, models of Lieblein, Koch-Smith, Aungier, Hawell are implemented to consider the COMPRESSOR losses. Finally, the model results are compared with experimental data to validate the model.

AXIAL slot CTs were designed and applied on Rotor 67 to understand the physical mechanisms responsible for the improvement of the stall margin. Unsteady Reynolds-averaged Navier-Stokes was applied in addition to steady Reynolds-averaged Navier-Stokes to simulate the flow field of the rotor. The results show that aerodynamic performance and the rotor stability were improved. Stall margin improvement (SMI) improved by 26. 85% after the CT covering 50% of the AXIAL tip chord was applied, whereas peak efficiency (PE) decreased the least. The main reason for the rotor stall in the solid casing is the blockage caused by tip leakage flow. After AXIAL slot CTs were applied, the tip leakage flow in the front part of the chord was obviously reduced, and the majority of the blockages in the tip region were removed. The absolute value of the AXIAL momentum before 45% AXIAL chord in CT_50 was reduced by 50%, whereas the maximum tangential momentum value of CT_50 was decreased by 70% relative to the solid casing. CT_50 configuration was located across the shock wave; thus, it can fully utilize the pressure gradient to bleed and remove the blockage region, and the across flow is considerably depressed.

The paper discusses the effect of COMPRESSOR characteristic on surge phenomena in AXIAL flow COMPRESSORs. Specifically, the effect of nonlinearities on the COMPRESSOR dynamics is analyzed. For this purpose, generalized multiple time scales method is used to parameterize equations in amplitude and frequency explicitly. The pure surge case of the famous Moore-Greitzer model is used as the basis of the study. The COMPRESSOR characteristic used in the Moore-Greitzer model is generalized to evaluate the effect of the parameters involved. Subsequently, bifurcation theory is used to study the effect of nonlinear dynamics on surge behavior. It has been found that the system exhibits supercritical Hopf bifurcation under specific conditions in which surge manifests as limit cycle oscillations. Key parameters have been identified in the analytical solution which govern the nonlinear dynamic behavior and are responsible for the existence of limit cycle oscillations. Numerical simulations of the Moore-Greitzer model are carried out and found to be in good agreement with the analytical solution

One of the factors that can cause a reduction in the efficiency and performance of AXIAL COMPRESSORs is the tip leakage flow of the COMPRESSOR blade. In the first, the COMPRESSOR's performance curve is compared with experimental results obtained under the condition of no air injection, and a statistically significant agreement is observed. The present study investigates the impact of various parameters, including flow rate, diameter, angle, and injection location, on the COMPRESSOR's performance curve and flow structure, taking into account the injection of air into a passage. The results indicate that the COMPRESSOR's stall margin and stable range extension are at their maximum values at a specific scale of each of the aforementioned parameters. Any deviation from this scale, either by reducing or increasing the injection parameters, leads to a reduction in the above characteristics. Although the presence of injection leads to an increase in the total pressure ratio in all injection states compared to the state without injection, the adiabatic efficiency at similar mass flow rates exhibits no significant change. The results also indicate that flow injection in the most suitable state increases the stall margin amount by 27% and the stable range extension of the COMPRESSOR by 192.

The correct materials selection in the design of aerospace structures reduces the weight and increases structural efficiency. Since the rotor in gas turbine engines has a significant weight, it is important to reduce its weight. The rotating disks in these rotors are subjected to mechanical and thermal loads and experience high-temperature gradients and angular velocities. This work aims to analyze the stress of a rotating disk made of carbon-carbon (C/C) composite to withstand mechanical and thermal loads and reduce the weight of the rotor. The behavior of constant thickness C/C composite disks is studied based on the Tsai-Wu Failure Theory. To do so, first, the basic properties of the material, disk size, rotation speed, temperature distribution, and other requirements are determined. The differential governing equations are obtained by assuming the plane stress state, Hooke's law, and compatibility condition, and the stresses, strains, and displacements are obtained. Considering the safety factors of 1 and 1.5, the critical velocities are calculated using the Tsai-Wu failure theory. Finally, according to the information obtained from the analysis, the evaluation of disks with different layers is compared with other similar disks made with different materials.