Advanced gas turbine cycles use advanced blade cooling technologies to reach high turbine inlet temperature. Accurate modeling and optimization of these cycles depend on blade cooling model. In this study, different models have been used to simulate gas turbine performance. The first model is the continuous model and the second is stage-by-stage model with alternative methods for calculating coolant, stagnation pressure loss and SPR. Variation of specific heat and enthalpy with temperature are included in both models. The composition of gas stream in turbine is changed step by step due to air cooling. These models are validated by two case study gas turbine results, which show good agreement with manufacture’s data. Then using these models, a comparison between continuous and stage by stage models is done. Results show that the stage by stage model can be improved to use in variable CPR by changing number of turbine stages. Also, the number of film cooled rows and coated rows change the number of cooled rows. This also increases power, efficiency and TET, while reduces coolant mass flow. For high TIT and with current blade cooling technology, efficiency seems to have no increase relative to lower TIT. CAP increases power but decreases efficiency, where with FP efficiency also increases, while power increases too.
This paper describes a new quasi-3D design method for centrifugal compressor impeller.The method links up a novel inverse design algorithm, called Ball-Spine Algorithm (BSA), and a quasi-3D analysis.Euler equation is solved on the impeller meridional plane.The unknown boundaries (hub and shroud) of numerical domain are iteratively modified by BSA until a target pressure distribution in flow passage is reached.To validate the quasi-3D analysis code, existing compressor impeller is investigated experimentally.Comparison between the quasi-3D analysis and the experimental results shows good agreement.Also, a full 3D Navier-Stokes code is used to analyze the existing and designed compressor numerically.The results show that the momentum decrease near the shroud wall in the existing compressor is removed by hub-shroud modifications resulting an improvement in performance by 0.6 percent.
In this research a multilevel optimization on the profile of splitter blades of a turbocharger compressor is performed using genetic algorithm in order to improve its performance. Successive corrections of profile at hub, midspan and shroud of splitter blades, with the objective of decreasing incidence losses at the leading edge and adjustment of blade loading at shroud, results in an impeller having improved splitter blades. The impeller flow filed analysis shows the optimization has been successful in reducing flow leakage at the shroud region and as well as losses in leading edge region. Although numerical simulations predict0.5% decrease in pressure ratio at design point, but 2.2 points improvement in isentropic efficiency is calculated. Based on the optimization results a new impeller is designed and manufactured and tested on a turbocharger test bed. Experimental results approve the simulation prediction results on the expected improvement in performance
Design and optimization of centrifugal compressors, based on main blades configuration of impeller have been vastly discussed in open literature, but less researches have addressed splitters. In this research, the impeller of a commercial turbocharger compressor is investigated. Here, profiles of main blades are not changed while the effect of changing the configuration of splitters is studied. An optimization study is performed to find the best configuration using genetic algorithm over a complete operating curve of the compressor. CFD codes with experimental support are used to predict the compressor performance. Quantumetric tests beside destructive analysis of two impellers are implemented for material identification and selection which is necessary for manufacturing. After taking into account structural considerations and approving the safety by numerical simulation, the new impeller is manufactured using 5 axis CNC machine. Non destructive tests are performed for identification of any structural defects. The new impeller is then mounted on a turbocharger shaft and tested experimentally in a wide range of operating conditions, which leads to a design having 2.3% improvement in efficiency. This is an important achievement in all applications of centrifugal compressors, especially in turbochargers.
In this research, design methods of radial flow compressor volutes are reviewed; the main criteria in volute primary designs are recognized and the most effective ones are selected. The effective parameters, i.e., spiral cross-section area, circumferential area distribution, exit cone, and tongue area of the compressor volute are parametrically studied to identify the optimum values. A numerical model has been prepared and verified through experimental data which are obtained from the designed turbocharger test rig. Different volutes are modeled and numerically evaluated using the same impeller and vane-less diffuser. For each model, the volute total pressure ratio, static pressure recovery and total pressure loss coefficients and the radial force on the impeller are calculated for different mass flow rates at design point and off-design conditions. The volute which shows better performance and causes lower the net radial force on the impeller at desired mass flow rates is selected as an optimal one. The results show the volute design approach differences at the design point and off-design conditions. Improving the pressure ratio and reducing total pressure loss at design point may result in the worse conditions at off-design conditions as well as increasing radial force on the impeller.
In this study, the main objective is to develop a one dimensional model to predict design and off design performance of an operational axial flow compressor by considering the whole gas turbine assembly. The design and off-design performance of a single stage axial compressor are predicted through 1D and 3D modeling. In one dimensional model the mass, momentum and energy conservation equations and ideal gas equation of state are solved in mean line at three axial stations including rotor inlet, rotor outlet and stator outlet. The total to total efficiency and pressure ratio are forecasted using the compressor geometry, inlet stagnation temperature and stagnation pressure, the mass flow rate and the rotational speed of the rotor, and the available empirical correlation predicting the losses. By changing the mass flow rate while the rotational speed is fixed, characteristic curves of the compressor are obtained. The 3D modeling is accomplished with CFD method to verify one dimensional code at non-running line conditions. By defining the three-dimensional geometry of the compressor and the boundary conditions coinciding with one dimensional model for the numerical solver, axial compressor behavior is predicted for various mass flow rates in different rotational speeds. Experimental data are obtained from tests of the axial compressor of a gas turbine engine in Sharif University gas turbine laboratory and consequently the running line is attained. As a result, the two important extremities of compressor performance including surge and choking conditions are obtained through 1D and 3D modeling. Moreover, by comparing the results of one-dimensional and three-dimensional models with experimental results, good agreement is observed. The maximum differences of pressure ratio and isentropic efficiency of one dimensional modeling with experimental results are 2.1 and 3.4 percent, respectively.
The axial turbine is one of the most challenging components of gas turbines for industrial and aerospace applications. With the ever-increasing requirement for high-aerodynamic performance blades, three-dimensional aerodynamic shape optimization is of great importance. In this research, the rear part of a gas turbine consisting of a one-stage axial turbine is optimized numerically. A useful optimization algorithm is presented to improve the efficiency and/or pressure ratio of the axial turbine with two different objective functions. The three-dimensional blade-shape optimization is employed to study the effects of the turbine stator and rotor lean and sweep angles on the turbine performance. The investigation is carried out at the turbine design speed. By coupling a verified computational fluid dynamics simulation code with the genetic algorithm, an automated design procedure is prepared. Geometry candidates for the optimization algorithm are generated by re-stacking of the two-dimensional airfoil sections. Three-dimensional, turbulent, and compressible flow field is numerically investigated via a Navier–Stokes solver to calculate various objective functions. Experimental results of the gas turbine are used for specifying the boundary conditions and validation of the simulation results. The proposed method results in 1.3% and 1.5% improvements in the turbine stage efficiency in design speed and reduced mass parameter at choke condition, respectively.
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