Guotai Feng

Effect of Turning Angle on Flow Field Performance of Linear Bowed Stator in Compressor at Low Mach Number

Abstract A comparison of the results of a computational simulation and an experimental measurement indicates a good agreement between them: the bowed blade lowers the energy loss coefficient of engine by 11% in the simulation and by 13% in the measurement. To further discuss the application conditions of bowed blade in compressor, with incidence equal to zero and other boundary conditions unchanged, a computational investigations on four series of linear stators with different aerofoil turning angles are achieved. It is found that the bowed blade has much positive effect in high airfoil turning angle cascade, for example, the optimal retrofit of 30° bow angle highly reduces the energy loss coefficient by 17.9%, when the aerofoil turning angle is 59.5°. But the optimal retrofit of 15° has only 0.7% reduction when the aerofoil turning angle is 39.5°, or even the compressor performance will get worse with the bow angle gradually increasing. Consequently, it is verified that the turning angle is one of the important factors to decide whether to apply the bowed blade into compressor at low Mach number.

Full Three-Dimensional Optimization Platform of Turbine Blades Considering the Film Cooling

In this paper, an optimization platform was established with Isight, cfx and the self-programming program which is used to generate the mesh. Film cooling effect can be taken into account. 15 parameters are selected as optimization variables. During the optimization process, the baseline blade and cooling holes are given by parameterized method. There are two objective functions during the optimization process. The first one is aerodynamic efficiency and the second one is film cooling efficiency. As there are two objective functions, NSGA-II is chosen as the multi-objective optimization algorithm. Then the Pareto-optimal front can be got. The results show that aerodynamic efficiency and film cooling efficiency restrict each other. It’s impossible to get the best solutions in one example, so the Pareto optimal set can provide a lot of choices. Different shapes make different effects on the aerodynamic efficiency and film cooling efficiency. From the above, it can be seen that the platform is helpful especially in the case that aerodynamic efficiency and film cooling efficiency restrict each other. This paper also discusses the prospects for platform applications.© 2013 ASME

Highly-Loaded Low-Reaction Boundary Layer Suction Axial Flow Compressor

In order to design high pressure ratio and highly loaded axial flow compressor, a new design concept based on Highly-Loaded Low-Reaction and boundary layer suction was proposed in this paper. Then the concept’s characteristics were pointed out by comparing with the MIT’s boundary layer suction compressor. Also the application area of this design concept and its key technic were given out in this paper. Two applications were carried out in order to demonstrate the concept. The first application was to redesign a low speed duplication-stage axial flow compressor into a single stage. The second one was a feasibility analysis to decrease an 11 stage axial compressor’s stage count to 7 while not changing its aerodynamic performance. The analysis result showed that the new design concept is feasible and it can be used on high pressure stage of the aero-engine, compressor of ground gas turbine (except the transonic stage) and high total pressure ratio blower.Copyright

Numerical Simulation of 3D Flow Field Structure in Turbine Cascade With Bowed Blades

The differences of flow field in bowed blade cascade and that in straight blade cascade are systematically studied in this paper. To bow a blade means to change its geometric boundary condition. This change not only affect the pressure distribution along the blade profile exit Mach number but also has great effect on the original position and development of the passage vertex. All of the changes mentioned above have great influence on the loss.Numerical simulation result showed that blade bowing can decrease the cross-pressure gradient near the end wall. This trend will be more obvious with the increase of the bow angle. The pressure gradient decrease is beneficial to weaken the passage vortex strength and reduces the secondary loss near the endwalls. In addition, Pressure gradient from endwalls to midspan can be established near suction surface in positively bowed blade. With the increase of bow angle, this C-type pressure distribution is remarkable. It is also found that this C-type pressure distribution will influence the position of corner vortex near the suction surface and will also influence the position and size of the passage vortex. Blade bowing also has great influence on the position of the saddle point near the leading edge and the separated line of the horseshoe vortex. It is found that the position of the saddle point and the separated line of both legs of the horseshoe vortex move forward in a positively bowed blade.The passage vortex structure in bowed cascade is also presented. It can be concluded that a bowed blade can make the passage vortex stable and helps change its structure from loose to compact. Blade bowing is also beneficial to limit the influence domain of the unstable passage vortex core by the stable limit cycle.Copyright

Study on Numerical Methods for Conjugate Heat Transfer Simulation of an Air Cooling Turbine

The effects of several numerical methods, including computational grids, coupling method, transition model and inner cooling air flow prediction, on the conjugate simulations were studied in the research. Firstly a finite difference conjugate solver was developed. Such solver included an N-S solver and a thermal conduction module for fluid flow and solid thermal conduction, respectively. Then conjugate simulations of an air cooling turbine were carried out. There were four kinds of conjugate simulations: the first one employs different types of computational grids, including H-type grids and O-type grids, for discretizing near-wall regions in fluid zone; the second one employs different coupling methods including indirect and direct ones; the third one employs different models including the B-L and q-ω turbulence models, and the AGS transition model; and the forth one employs different turbulence models for the prediction of flows in the cooling channels. All of the numerical results have been compared to the experimental result. Finally it concludes that to accurately predict thermal and aerodynamic load of the air cooled turbine, the conjugate simulation should employ O-type girds to discretize the near wall regions in the fluid zone, use the direct coupling method to transfer data between solid and fluid domains, and utilize the transition model to predict accurate flow details within the boundary layers, and also account for flows in the cooling air channels.© 2008 ASME

Calculation of the Energy Loss for Tip Leakage Flow in Turbines

Abstract A commercial N-S solver has been employed for simulation and investigation of the unsteady flow field inside the tip clearance of a turbine rotor. The main objective of this paper is to introduce a new method of energy loss calculation for the flow field in tip clearance region of a turbine rotor blade This method can be easily used in all kinds of flow fields. Regions of high viscous effects have been found to be located near the shroud rather than the blade tip. It is shown that the time-averaged loss of energy in tip leakage flow is dissimilar for different rotor blades. This result is a helpful hint that can be taken by blade designers to design non uniform rotor blades with different geometric and aerodynamic loads to minimize the energy loss.

The Development of Highly Loaded Turbine Rotating Blades by Using 3D Optimization Design Method of Turbomachinery Blades Based on Artificial Neural Network & Genetic Algorithm

In order to improve turbine internal efficiency and lower manufacturing cost, a new highly loaded rotating blade has been developed. The 3D optimization design method based on artificial neural network and genetic algorithm is adopted to construct the blade shape. The blade is stacked by the center of gravity in radial direction with five sections. For each blade section, independent suction and pressure sides are constructed from the camber line using Bezier curves. Three-dimensional flow analysis is carried out to verify the performance of the new blade. It is found that the new blade has improved the blade performance by 0.5%. Consequently, it is verified that the new blade is effective to improve the turbine internal efficiency and to lower the turbine weight and manufacturing cost by reducing the blade number by about 15%.

Impact of Different Film-Cooling Modes at Leading Edge on the Aerodynamic and Heat Transfer Performance of Heavy Duty Gas Turbine

In this paper, the leading edge film-cooling flow field of a heavy duty gas turbine cascade has been studied by central difference scheme and multi-block grid technique. The research is based on the three-dimensional N-S equation solver. By way of comparison and analysis of the temperature field, the distribution of profile pressure, and the distribution of film-cooling adiabatic effectiveness in the region of leading edge with different cool air mass and injection angles, it is found that the aerodynamic energy loss drops a little by adding the cool air, the distribution of temperature of the blade is obviously changed and the adiabatic effectiveness at the leading edge and suction side is higher than that on pressure side. Profile pressure is not changed obviously in the whole, with the exception in the local region near the cooling holes. The change of the pressure variation is greater on the suction side. The influence of the change of cool air mass and injection angles on the flow field near the leading edge is obviously.

Numerical Study on End-Wall Flow in Highly Loaded Supercritical Compressor Cascades

This paper presents a numerical study on three-dimensional flow phenomena near the endwall of a linear high-turning compressor cascade at supercritical flow conditions. The compressor cascade with 60° camber angle was designed at a higher supercritical speed (M1 >0.9) by optimum method based on the baseline which aimed at improving the flow near the stator hub of small transonic fans. The camber line and thickness distribution curves of the baseline are formed by quadratic polynomials and double cubic curves respectively. The stack line and the thickness distribution near the end-wall were chosen as optimization variables to approach the objective function of total pressure loss coefficient, since they are the two main geometry parameters which can influence end-wall flow obviously. The analysis in current paper focuses on comparing the flow phenomena near the end-wall of baseline cascade with that of optimized one. Numerical simulation results are presented to show the loss reduction from the baseline to the optimized cascade near end-wall. The boundary-layer development on the suction surface, flow separation structure, shockwave and local supersonic area on the suction surface near the end-wall are analyzed in detail. The optimized cascade has a stronger shockwave near the leading edge. It was found that the radial flow of the boundary-layer caused by the optimization of stack line is the key factor influencing the aerodynamics loss near the end-wall at supercritical condition which also plays an important part in second-flow and flow separation in the corner. An understanding of the low-loss pattern of the end-wall flow and the flow filed structure for high-turning compressor at higher supercritical flow conditions then is summarized at the end of this paper.Copyright

The Key Techniques for Thermal-Flow-Elastic Coupling Numerical Simulation Platform in Turbines

Thermal-flow-elastic coupling (TFEC) numerical simulation platform has been an essential platform for designing turbo engines with high performance and efficiency. Generally, TFEC numerical simulation was achieved by predicting thermal and stress fields with finite element methods, while flow fields with finite difference methods, but such calculation was not popular in engineering design, because of too much size of data exchange and lower computing efficiency. However these shortcomings will not exist by using finite difference methods for all of the fields. To establish a three dimensional multifunction numerical simulation platform for turbines for all of the fields, the key technique was studied firstly. The technique included analysis on physical models, establishing of mathematical model equation, usage of curvilinear coordinate platform, construction of high accuracy difference scheme and selection of boundary conditions for multi-field coupling simulation. Then the algorithm including domain decomposition one and parallel one were studied to accelerate the coupling simulation. The purpose was to develop a completed TFEC numerical simulation platform by using of finite difference method and to apply the platform for numerical simulation in turbines. Firstly codes for predicting flow field in passage, thermal and stress fields in solid body were developed. Then a simple TFEC numerical simulation platform for turbines was obtained. The single code for predicting flow field was verified with experimental data, and the other two codes were validated with thermal and elastic analytic solutions respectively. And satisfying results were obtained. Then the code for thermal-flow was validated with experimental data of Mark II blade, and the code for thermal-elastic coupling simulations was validated with a cylinder by an analytic solutions. All of these are good basics for completing TFEC numerical simulation platform using finite difference methods for all of the fields and computing TFEC numerical simulation in a turbo engine.Copyright