Songtao Wang

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

A New Design Concept of Highly-Loaded Axial Flow Compressor by Applying Boundary Layer Suction and 3D Blade Technique

A new design concept of highly-loaded axial flow compressor by applying boundary layer suction and 3D blade technique was proposed in this paper. The basic idea of this design concept was that low reaction was adopted as while as increasing the rotor’s geometry turning angle, so that the boundary layer separation of a rotor could be eliminated and the rotor was kept working in high efficiency. This design concept would greatly increase the stator’s geometry turning angle, so boundary layer suction on stator cascades was adopted in order to restrain the boundary layer separation. In some situations, 3D blade technique was also applied in order to control the boundary layer separation more efficiently. The advantages of the above design concept were: the compressor’s pressure ratio was increased remarkably; boundary layer suction was only adopted in stator cascades so as to reduce the complexity of boundary layer suction structure. The key techniques of the new design concept were also explained in this paper. In order to increase the compressor’s pressure ratio, the geometry turning angle of rotor was increased greatly, and the rotor inlet was prewhirled to reduce the rotor’s reaction so as to restrain the rotor’s boundary separation. Boundary layer suction was carried out in the stator cascades (mainly on suction side), hub and shroud in order to control the flow separation. 3D blade technique could be adopted if necessary. The limitation of the application of this design concept was also pointed out through the analysis of the Mach number at rotor inlet, the prewhirl angle of rotor, the work distribution along span wise and the control method of stator separation. Numerical simulation was carried out on a single low-reaction compressor stage with IGV in order to demonstrate the new design concept. By using boundary layer suction and 3D blade technique, the energy loss in stator cascades was greatly reduced and the whole stage’s isentropic efficiency was about 90%. The low-reaction stage’s aerodynamic load was double than conventional design. The boundary layer separation could be effectively controlled by proper combination of boundary layer suction and bowed or twisted blade. The numerical result proved that the new design concept was feasible and had a wide application area.Copyright

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.

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 Numerical Study of a New Air-Cooled Turbine Blade With Depression Feature at Leading Edge

A new air-cooled turbine blade with depression feature at leading edge was presented in this paper. The flow field of this blade at the transonic condition was simulated in detail. The change of vortex position induced by the leading edge cooling air injection along the blade profile was observed. The aerodynamic characteristics of the new blade such as film effectiveness, energy loss, pressure field etc. were compared with the traditional air-cooled turbine blade. The result shows that this new blade locks the vortex induced by the leading edge cooling air injection in its depression part. The mixing of cooling air and mainstream becomes weak. Thus, the number of the leading edge cooling air injection holes can be reduced. Correspondingly, the amount of injection can also be reduced.Copyright